Supersonic compressor

ABSTRACT

A supersonic compressor including a rotor to deliver a gas at supersonic conditions to a diffuser. The diffuser includes a plurality of aerodynamic ducts that have converging and diverging portions, for deceleration of gas to subsonic conditions and then for expansion of subsonic gas, to change kinetic energy of the gas to static pressure. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when the aerodynamic ducts are designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of in excess of two to one, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct. In an embodiment, the number of leading edges are minimized, and may be less than half, compared to the number of blades in the accompanying rotor.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority from prior pending U.S. ProvisionalPatent Application Ser. No. 61/506,055, for a SUPERSONIC COMPRESSOR,which was filed on Jul. 9, 2011, and which is incorporated herein in itsentirety by this reference.

STATEMENT OF GOVERNMENT INTEREST

This invention was made with United States Government support underContract No. DE-FE0000493 awarded by the United States Department ofEnergy. The United States Government has certain rights in theinvention.

COPYRIGHT RIGHTS IN THE DRAWING

A portion of the disclosure of this patent document contains materialthat is subject to copyright protection. The applicant has no objectionto the facsimile reproduction by anyone of the patent document or thepatent disclosure, as it appears in the Patent and Trademark Officepatent file or records, but otherwise reserves all copyright rightswhatsoever.

TECHNICAL FIELD

This description relates to apparatus and methods for the compression ofgases, and more particularly to gas compressors which are designed toutilize supersonic shock compression.

BACKGROUND

A continuing interest exists in industry for a simple, highly efficientgas compressor. Such devices may be useful in a variety of applications.Operational costs could be substantially improved in many applicationsby adoption of a compressor that provides improvements in operatingefficiency as compared to prior art compressor designs. Further, fromthe point of view of maintenance costs, it would be desirable to developnew compressor designs that reduce the mass of rotating components,since rotating components have generally been identified ascomparatively costly when replacement or repair becomes necessary, ascompared to non-rotating parts which are subject to stress and strainfrom temperature and pressure, but not to additional loads due to rotarymotion. Thus, it can be appreciated that it would be advantageous toprovide a new, high efficiency compressor design which minimizes movingparts.

Although a variety of supersonic compressors have been contemplated, andsome have been tested by others, the work of J. K. Koffel et al. asreflected in U.S. Pat. No. 2,974,858, issued Mar. 14, 1961, and entitled“High Pressure Ratio Axial Flow Supersonic Compressor,” the disclosureof which is incorporated herein in its entirety by this reference, isinstructive of such work generally, and thus is suggestive of technicalproblems that remain in the field and with respect to which bettersolutions are required in order to improve operational capability andcompression efficiency. Although the Koffel et al. patent describes theuse of an impulse blade rotor and illustrates a downstream bladedstator, the compressor geometry described would appear, maximally, toonly enable achievement of pressure ratios stated therein, which are atone point mentioned as an “ . . . overall pressure ratio ofapproximately 4 to 1 in a single rotor-stator stage.” And, although theKoffel et al. patent mentions issues with respect to boundary layereffects, it does not provide for integrated control of such phenomenonas may be useful to avoid perturbations caused by boundary layerinteraction with shock waves, especially as might be applied forcompressor operation at higher pressure ratios than those noted therein.

In short, there remains a need to provide a design for a high pressureratio supersonic compressor that simultaneously resolves variouspractical problems, including (a) providing for starting of a compressordesigned for high pressure ratio operation so as to control a normalshock at an effective location in a supersonic diffuser designed forhigh pressure ratio and efficient compression, (b) avoiding excessivenumbers of leading edge structures (such as may be encountered in priorart multi-bladed stators), and minimizing other losses encountered by ahigh velocity supersonic gas flow stream upon entering a diffuser, and(c) providing for effective boundary layer control, especially asrelated to retention of a normal shock at a desirable location, in orderto achieve high compression ratios in an efficient manner.

SUMMARY

A novel supersonic compressor has been developed that, in an embodiment,minimizes the number of rotating parts. The compressor utilizes a rotorhaving a plurality of blades extending into a gas flow passage todevelop gas velocity in an incoming gas flow stream, and to acceleratethe incoming gas flow stream tangentially and axially, to deliver a gasflow stream at supersonic conditions to a diffuser that includes one ormore aerodynamic ducts. In an embodiment, a plurality of blades areprovided as impulse blades, in that they provide kinetic energy toincrease gas velocity to supersonic conditions, with little if anystatic pressure rise. In an embodiment, the number of aerodynamic ductsis minimized. As a result, a small number of inlets (at least one inletbeing associated with each aerodynamic duct) may be utilized, ratherthan a large number of stator blades. In an embodiment, an exemplarydesign minimizes the total number of leading edges, and thus the lengthof leading edges exposed to the incoming supersonic gas flow isminimized. In an embodiment, the aerodynamic ducts of the diffuser maybe wrapped about a surface of revolution that extends along alongitudinal axis, for example, on a cylindrical shape or partialconical shape. In an embodiment, aerodynamic ducts may be provided in ahelical or helicoidal configuration. In an embodiment, the aerodynamicducts may be provided in a shape having a relatively constant helicalangle. In an embodiment, the aerodynamic ducts may be provided along acenterline in the general configuration of a circular helix, in that theratio of curvature to torsion is constant. Other helical shapes may beprovided, including shapes with differing ratios of curvature totorsion. Without limitation, various examples are provided herein. Forexample, in an embodiment, aerodynamic ducts may be provided in a conichelix configuration, in the form of a slight spiral as if located overan underlying conic surface. In various embodiments, aerodynamic ductsmay be either right handed or left handed, with inlets and throatsoriented substantially with the direction of high pressure supersonicgas leaving the blades of a rotor. Other embodiments may utilize othershapes (for example, non-helical or other shapes) for aerodynamic ducts,and thus the suggested shapes described herein are merely forexplanation, without limitation thereby. A series of oblique shocks anda normal shock may be utilized within the aerodynamic ducts toefficiently transform the high velocity incoming supersonic gas flow toa high pressure subsonic gas flow. Subsequent to a first stationarydiffuser, gas velocity may be further reduced and static pressure may beaccumulated by volute or other suitable structure known in the art.Alternately, a second compression stage may be utilized. In anembodiment, a second compression stage may accept as inlet gas thecompressed gas output from a first compression stage. The secondcompression stage may have a second rotor with a plurality of bladesextending into a gas flow passage, and a second stator including furtheraerodynamic ducts, in order to further compress gas after it leaves afirst compression stage. And further stages of compression (e.g., inexcess of two stages), may be utilized for yet higher overallcompression ratios for particular applications.

For starting a system of supersonic shock waves, in an embodiment, adiffuser may include bypass gas outlets for removal of a portion of theincoming gas flow to an extent that facilitates the establishment ofsupersonic shocks within the diffuser, consistent with a design pointfor a selected compression ratio, inlet Mach number, and mass flow of aselected gas. In an embodiment, the bypass gas outlets may be utilizedfor recycle of a portion of incoming gas, for passage through blades ofthe rotor, and thence back to an inlet for the aerodynamic ducts. In anembodiment, particularly for compression of air, the bypass gas may besimply discharged to the atmosphere. In an embodiment, the gascompressor may provide geometrically adjustable portions in aerodynamicducts, to change the quantity of incoming gas flow through the diffuser,in order to start and establish stable supersonic shock operation. In anembodiment, both starting bypass gas outlets and geometricallyadjustable portions may be utilized.

For minimization of adverse aerodynamic effects, and for improvingefficiency of gas flow through a diffuser, one or more boundary layercontrol structures may be utilized. Such boundary layer controlstructures may be selected from one or more types of boundary layercontrol techniques, including removal of a portion of gas flow viaboundary layer extraction or bleed, or by energizing a boundary layer byboundary layer gas injection, or by energizing a boundary layer bymixing, such as by use of vortex generators. In an embodiment, thevortex generators may generate multiple vortices, wherein a largervortex rotates a simultaneously generated, adjacent, and smaller vortextoward and thence into a boundary layer, and thus controls such boundarylayer as the smaller vortex mixes with the boundary layer.

In an embodiment, the compressor described herein may have multiple gaspaths, that is, multiple aerodynamic ducts, for generating supersonicshock waves and for allowing subsonic diffusion downstream of a throatportion. Since, in an embodiment, supersonic shocks may be locatedwithin stationary structures, such as along a stationary ramp portion ofan aerodynamic duct, the control of shock location is greatlysimplified, as compared to various prior art supersonic compressordesigns where shocks are located between structures in rotors, orbetween rotors and adjacent stationary structures such ascircumferential walls.

Further, the location of shocks within stationary diffusers avoidsparasitic losses that are present in prior art designs due to dragresulting from the rotational movement of various rotor components. Morefundamentally, an embodiment of the compressor design disclosed hereindevelops high compression ratios with very few aerodynamic leading edgestructures, particularly stationary structures, protruding into thesupersonic flow path. In part, such improvement is achieved because adesign is provided in which the number of aerodynamic ducts isminimized. In an embodiment, only a single leading edge is provided peraerodynamic duct, and thus the number of leading edge surfacesinterposed into a supersonic gas flow stream is minimized. Consequently,the compressor design(s) disclosed herein have the potential to providehighly efficient gas compressors, as compared to heretofore known gascompressors, especially when operating at high compression ratios in asingle compression stage. For example, and without limitation, thecompressor designs disclosed herein may operate at compression ratios ina single stage of up to about four to one (4:1), or at least about fourto one (4:1), or at least about six to one (6:1), or from between aboutsix to one to about ten to one (about 6:1 to about 10:1), or up to abouttwelve and one-half to one (12.5:1), or higher than twelve to one(12:1).

Finally, many variations in gas flow configurations, particularly indetailed rotor geometry and in detailed diffuser geometry, may be madeby those skilled in the art and to whom this specification is directed,without departing from the teachings hereof.

BRIEF DESCRIPTION OF THE DRAWINGS

Configurations for novel supersonic compressors will be described by wayof exemplary embodiments, using for illustration the accompanyingdrawing figures in which like reference numerals denote like elements,and in which:

FIG. 1 is a partially cut-away vertical view, showing, in cross-section,an inlet passageway feeding a gas supply to impulse blades on a rotor(shown from the side to reveal exposed blades). The impulse bladesdeliver gas at supersonic conditions to a stationary diffuser having aplurality of aerodynamic ducts. The aerodynamic ducts include convergingand diverging portions, inlet bypass gas passageways for starting, andboundary layer outlet bleed ports for boundary layer control. FIG. 1also shows an embodiment for a diffuser in which the throat of theaerodynamic duct is in close alignment with the direction of gas flowleaving the rotor blades.

FIG. 2 provides a gas velocity diagram with respect to an exemplaryimpulse rotor blade design, describing gas velocity components at fourdifferent locations relative to blades extending from a rotor.

FIG. 3 is a perspective view of impulse blades on a rotor and adownstream stationary diffuser that includes a plurality of aerodynamicducts, showing a helical structure for the aerodynamic ducts havingconverging and diverging portions, as well as inlet bypass passagewaysfor starting, and boundary layer ports for boundary layer control, andportions of adjacent static structure in phantom lines.

FIG. 4 is a partial cross-sectional perspective view of an embodimentfor a compressor, showing an inlet passageway, impulse blades on arotor, a stationary diffuser including an aerodynamic duct havingconverging and diverging portions, and boundary layer bleed passageways.

FIG. 5 is a cross-sectional view of a stationary diffuser including theuse of five (5) aerodynamic ducts having converging and divergingportions, as well as inlet bypass passageways for starting, and boundarylayer bleed ports for boundary layer control, as well as associatedsub-chambers and passageways adjacent the converging and divergingportions.

FIG. 6 is an enlarged detail of a portion of an exemplary aerodynamicduct similar to that first depicted in FIG. 5, but now showing the use,in an embodiment, of boundary layer bleed through outlet bleed ports forboundary layer control, and at the same time, the use of vortexgenerators within the aerodynamic duct for control of a boundary layerby mixing.

FIG. 7 provides a circumferential view of an exemplary gas flow pathinto an impulse bladed rotor and thence through a diffuser havingleading edges followed by a plurality of aerodynamic ducts each having aconverging portion provided via a compression ramp and a divergingportion illustrated by expansion ramps, and showing bypass bleedpassageways for starting, and boundary layer outlet bleed ports toassist in boundary layer control, for shock stability, and forefficiency.

FIG. 8 is an enlargement of a portion of the circumferential view justprovided in FIG. 7, now showing a leading edge of an aerodynamic duct ina diffuser, and also showing a converging portion provided via acompression ramp and diverging portion illustrated by an expansion ramp,and showing starting bleed ports and boundary layer ports.

FIG. 8A is an enlarged portion of FIG. 8, showing a leading edge wedgeangle for a stator, and a partition wall located rearward, i.e.downstream therefrom which, in an embodiment, may be configured as acommon partition to separate adjacent aerodynamic ducts in a stationarydiffuser.

FIG. 8B is cross-section taken across line 8B-8B of FIG. 8A, showing aleading edge for an aerodynamic duct, and more specifically, how aleading edge may, in an embodiment, be provided in a swept-backconfiguration, that is sloping rearward in the flowwise direction.

FIG. 8C is an enlarged portion of FIG. 8A, showing a suitable radius fora leading edge of an aerodynamic duct.

FIG. 9 is a vertical cross-section of an embodiment for a compressor,showing a gas passageway for incoming gas to be compressed, and adiffuser including a stationary aerodynamic duct with converging anddiverging portions, and a volute for deceleration of gas andaccumulation of static pressure, as well as an associated gearbox andmotor.

FIG. 10 is a perspective view of an embodiment for an impulse rotor,similar to that seen in FIG. 3 above, but now showing the use of animpulse rotor having a shroud for the blades, and in this embodiment,also showing teeth for a labyrinth-type seal structure on thecircumferential portions of the rotor shroud.

FIG. 11 is a partial cross-sectional view of an embodiment for acompressor, similar to that shown in FIG. 4 above, showing an inletduct, impulse rotor having a shroud such as just illustrated in FIG. 10,a diffuser including an aerodynamic duct having geometrically adjustableconverging and diverging portions and which is adapted for changing theeffective contraction ratio of the aerodynamic duct for starting andsetting up a supersonic shock wave in a suitable location, and furthershowing the use of vortex generators for effective control of boundarylayer phenomenon.

FIG. 12 is a schematic cross-sectional view of an embodiment foradjustable converging and diverging portions located within anaerodynamic duct as first illustrated in FIG. 11 above, now furthershowing how adjustment of the duct changes the effective contractionratio (also known as convergence ratio) in the duct by adjusting thearea available for passage of gas therethrough.

FIG. 13 is a schematic cross-sectional view of an embodiment for anaerodynamic duct including converging and diverging portions, includinga stationary diffuser, illustrating both the use of a gas removal andbypass system for starting, and use of a boundary layer bleed system forcontrol of boundary layer phenomenon.

FIG. 13A is a partial cross-sectional view of an embodiment for anaerodynamic duct including converging and diverging portions,illustrating both the use of an openable door for gas removal duringstarting, and the use of boundary layer bleed systems for control ofboundary layer phenomenon.

FIG. 14 is a partial cross-sectional perspective view of an embodimentfor a compressor, similar to that shown in FIGS. 3 and 11 above, showingan inlet duct, impulse blades with shroud on a rotor, a diffuserincluding an aerodynamic duct utilizing a gas removal system forstarting of the type just set forth in FIG. 13 above, and furthershowing the use of a boundary layer bleed system for effective controlof boundary layer phenomenon.

FIG. 15 is a vertical cross-sectional view taken at line 15-15 of FIG.1, showing an embodiment for an entrance to a diffuser, here showingfive (5) aerodynamic ducts, and further showing short height of leadingedges of the aerodynamic ducts.

FIG. 16 is a vertical cross-sectional view taken as if at line 16-16 ofFIG. 1, but now showing the entrance to an alternate embodiment using adiffuser having seven (7) aerodynamic ducts, and further showing a shortheight for leading edges of the aerodynamic ducts.

FIG. 17 is a diagrammatic side view for an embodiment for a compressor,depicting the use of an impulse bladed rotor (possible additional bladeshroud is not shown) with a diffuser including a plurality ofaerodynamic ducts located around a surface of rotation, in an embodimenthelicoidally, and wherein the surface of rotation as indicated by brokenlines is generally cylindrical in shape.

FIG. 18 is a diagrammatic side view for an embodiment for a compressor,depicting the use of an impulse bladed rotor (possible additional rotorshroud is not shown) with a diffuser including a plurality ofaerodynamic ducts located around a surface of rotation, in an embodimentin a generally spiral configuration, and wherein the surface of rotationas indicated by broken lines is generally in the shape of an outwardlysloping truncated cone.

FIG. 19 is a diagrammatic side view for an embodiment for a compressor,depicting the use of an impulse bladed rotor (possible additional shroudis not shown) with a diffuser including aerodynamic ducts located arounda surface of rotation, in an embodiment in a generally spiralconfiguration, and wherein the surface of rotation as indicated bybroken lines is generally in the shape of an inwardly sloping truncatedcone.

FIG. 20 is a diagrammatic side view for an embodiment for a vortexgenerator affixed to a selected surface of an aerodynamic duct, whereinthe vortex is designed to generate at least one (1) vortex, and hereshowing the generation of two (2) vortices from an incoming gas flow asindicated by heavy broken lines.

FIG. 21 is a diagrammatic end view for the embodiment of a vortexgenerator as just illustrated in FIG. 20 above, showing two (2)vortices, a larger one and a smaller one, as first generated above aselected surface of an aerodynamic duct.

FIG. 22 is a diagrammatic end view for the embodiment of a vortexgenerator as just illustrated in FIGS. 20 and 21 above, showing two (2)vortices, a larger one and a smaller one, as the two vortices turn andflip the smaller vortex downward against the selected surface of anaerodynamic duct, so as to become located in a position for effectingwork on a boundary layer adjacent the selected surface.

FIG. 23 is a diagrammatic side view for an embodiment for a vortexgenerator affixed to a selected surface of an aerodynamic duct, whereinthe vortex is designed to generate at least one (1) vortex, and hereshowing the generation of three (3) vortices from an incoming gas flowas indicated by heavy broken lines.

FIG. 24 is a diagrammatic end view for the embodiment of a vortexgenerator as just illustrated in FIG. 23 above, showing three (3)vortices, a large one, an intermediate sized one, and a small one, asfirst generated above a selected surface of an aerodynamic duct.

FIG. 25 is a diagrammatic end view for the embodiment of a vortexgenerator as just illustrated in FIGS. 23 and 24 above, showing three(3) vortices, a large one, an intermediate sized one, and a small one,as they turn and flip the smaller vortices downward against the selectedsurface of an aerodynamic duct, so as to become located in a positionfor effecting work on a boundary layer adjacent the selected surface.

FIG. 26 is a partial cross-sectional view taken along the centerline ofan aerodynamic duct having a converging and diverging portion therein,showing the use of pressurized gas supplied by supply conduits for usein boundary layer control by gas injection.

FIG. 27 shows an enlarged portion of the partial cross-sectional viewprovided in FIG. 26, showing the use of a conduit for providing a supplyof gas for injection of a gas jet to control boundary layer buildup atthe wall near an expansion shock in an aerodynamic duct.

FIG. 28 is a cross-sectional view along the centerline of a generallyhelicoidal aerodynamic duct in a diffuser, showing an embodiment whereina compression ramp is located on an inward surface, and wherein bleedair passageways for starting are located on an outward surface of theaerodynamic duct, and also showing a plurality of oblique shockstructures S₁, S₂, S₃, and S_(x), as well as a normal shock S_(N), andthe use of vortex generators to control a boundary layer adjacent aradially interior surface of the aerodynamic duct.

FIG. 29 is a cross sectional view along the centerline of a generallyhelicoidal aerodynamic duct in a diffuser, wherein a compression ramp islocated on an outward surface, showing an embodiment wherein bypass gaspassageways for starting and establishing stable operation of the shockwave structure are located on an inward surface of the aerodynamic duct,and also showing a plurality of oblique shock structures S₁, S₂, S₃, andS_(x), as well as a normal shock S_(N), and the use of vortex generatorsto control a boundary layer adjacent an interior surface of theaerodynamic duct.

FIG. 30 is a cross-sectional view along the centerline of a generallyhelicoidal aerodynamic duct in a diffuser, wherein a compression ramp islocated on both an outward surface and on an inward surface, and showingan embodiment wherein bypass gas passageways for starting andestablishing stable operation of the shock wave structure are located onboth the outward surface and the inward surface of the aerodynamic duct,and also showing a plurality of oblique shock structures S₁, S₂, S₃, S₄,S₅, S₆, S₇, and S_(x), as well as a normal shock S_(N), and the use ofvortex generators to control a boundary layer adjacent an interiorsurface of the aerodynamic duct.

FIG. 31 is partial circumferential view showing the longitudinalcenterline of a diffuser, and the generally helical aerodynamic ductsused therein, as well as the accompanying rotor and its rotationalcenterline, showing an embodiment wherein a compression ramp is locatedon an outwardly extending trailing edge surface, and wherein bypass gaspassageways for starting and establishing stable operation of the shockwave(s) are located on the converging compression ramp surface, and alsoshowing a plurality of oblique shock structures S₁, S₂, S₃, and S_(x),as well as a normal shock S_(N).

FIG. 32 is partial circumferential view showing the longitudinalcenterline of a diffuser, and the generally helical aerodynamic ductsused therein, as well as the accompanying rotor and its rotationalcenterline, showing an embodiment wherein a compression ramp is locatedon an inward leading edge surface, and wherein bleed air passageways forstarting are located on the converging compression ramp surface, andalso showing a plurality of oblique shock structures S₁, S₂, S₃, andS_(X), as well as a normal shock S_(N).

FIG. 33 is partial circumferential view showing the longitudinalcenterline of a diffuser, and the generally helical aerodynamic ductsused therein, as well as the accompanying rotor and its rotationalcenterline, showing an embodiment wherein a compression ramp is locatedon an inward leading edge surface, and also on a trailing edge, andwherein bleed air passageways for starting are located on both of theconverging compression ramp surfaces, and also showing a plurality ofoblique shock structures S₁, S₂, S₃, S₄, etc., as well as a normal shockS_(N).

FIG. 34 is diagrammatic flow sheet depicting the use of at least twocompression stages, wherein the high pressure gas from a firstcompressor stage is provided to the low pressure entry of a secondcompressor stage for further compression.

FIG. 35 is a perspective view of an embodiment for an aerodynamic ductfor a diffuser in a compressor, showing the use of two vortex generatorsto aid in boundary layer control as gas flows through the aerodynamicduct.

FIG. 36 is a perspective view showing details of the vortex generatorsjust illustrated in FIG. 35, and now showing the height of vortexgenerators in an embodiment.

FIG. 37 is a plan view of the vortex generators just shown in FIGS. 35and 36, now illustrating the orientation of the vortex generators in adiffuser, where the vortex generators are offset by a selected anglefrom the centerline of the longitudinal axis of the aerodynamic duct ina diffuser.

FIG. 38 is a perspective view of an embodiment for an aerodynamic ductfor a diffuser in a compressor, showing the use of four vortexgenerators to aid in boundary layer control as gas flows through theaerodynamic duct.

FIG. 39 is a perspective view showing details of the vortex generatorsjust illustrated in FIG. 38, and now showing the height of vortexgenerators in an embodiment.

FIG. 40 is a plan view of the vortex generators just shown in FIGS. 38and 39, now illustrating the orientation of the vortex generators in adiffuser, where the vortex generators are offset by a selected anglefrom the centerline of the longitudinal axis of the aerodynamic duct ina diffuser.

The foregoing figures, being merely exemplary, contain various elementsthat may be present or omitted from actual supersonic compressor designsutilizing the principles taught herein, or that may be implemented invarious applications for such compressors. Other compressor designs mayuse slightly different aerodynamic structures, mechanical arrangements,or process flow configurations, and yet employ the principles describedherein or depicted in the drawing figures provided. An attempt has beenmade to draw the figures in a way that illustrates at least thoseelements that are significant for an understanding of an exemplarysupersonic compressor design. Such details should be useful forproviding an efficient supersonic compressor design for use inindustrial systems.

It should be understood that various features may be utilized in accordwith the teachings hereof, as may be useful in different embodiments asnecessary or useful for various gas compression applications, dependingupon the conditions of service, such as temperatures and pressures ofgas being processed, within the scope and coverage of the teachingherein as defined by the claims.

DETAILED DESCRIPTION

The following detailed description, and the accompanying figures of thedrawing to which it refers, are provided describing and illustratingsome examples and specific embodiments of various aspects of theinvention(s) set forth herein, and are not for the purpose ofexhaustively describing all possible embodiments and examples of variousaspects of the invention(s) described and claimed below. Thus, thisdetailed description does not and should not be construed in any way tolimit the scope of the invention(s) claimed in this or in any relatedapplication or resultant patent.

To facilitate the understanding of the subject matter disclosed herein,a number of terms, abbreviations or other shorthand nomenclature areused as set forth herein below. Such definitions are intended only tocomplement the usage common to those of skill in the art. Any term,abbreviation, or shorthand nomenclature not otherwise defined shall beunderstood to have the ordinary meaning as used by those skilledartisans contemporaneous with the first filing of this document.

In this disclosure, the term “aerodynamic” should be understood toinclude not only the handling of air, but also the handling of othergases within the compression and related equipment otherwise described.Thus, more broadly, the term “aerodynamic” should be considered hereinto include gas dynamic principles for gases other than air. For example,various relatively pure gases, or a variety of mixtures of gaseouselements and/or compounds, may be compressed using the apparatusdescribed, and thus as applicable the term “aerodynamic duct” shall alsoinclude the compression of gases or gas mixtures other than air, in whatmay be considered a gas dynamic duct.

The term “diffuser” may be used to describe an apparatus designed toreduce the velocity and increase the pressure of a gas entering atsupersonic velocity. A diffuser may employ one or more aerodynamicducts, which, when multiple aerodynamic ducts are used, divide theincoming gas into smaller flows for processing. Such aerodynamic ductsin a diffuser may include (a) a supersonic diffuser portion, which maybe in the form of a converging portion generally of decreasingcross-sectional area and which receives gas at supersonic velocity andcreates oblique shocks, (b) a throat portion, at or in which in aminimal throat cross-sectional area is provided, and (c) a subsonicdiffuser portion, which may be in the form of a diverging portion ofincreasing cross-section toward a final subsonic diffusercross-sectional area and which allows kinetic energy from gas velocityto be converted into static pressure of the gas.

The term “impulse blade(s)” may be used to describe blades used toaccelerate the flow of gas having a characteristic geometry whereinkinetic energy is imparted to the gas passing therethrough, and at atheoretical limit, no pressure increase is imparted to the gas passingtherethrough. Thus, in impulse blades as described herein, work done ona gas flow by impulse blades results predominantly in an increase invelocity, rather than predominantly in an increase in pressure. Thevelocity increase of a gas flow through impulse blades is achieved bychange of direction of the gas flow.

The term “inlet” may be used herein to define an opening designed forreceiving fluid flow, and more specifically, the flow of gas. Forexample, in an aerodynamic duct for a diffuser of a supersoniccompressor, the aerodynamic duct has an inlet having an inletcross-sectional area that is shaped to capture and ingest gas to becompressed. Inlets may have a large variety of shapes, and a fewexemplary shapes are provided herein.

The term “startup” may be used to define the process of initiating gasflow and achieving stable supersonic flow of gas through a convergingportion, and into at least some of a diverging portion of generallyincreasing cross-sectional area extending downstream from a throat of anaerodynamic duct. More specifically, startup is the achievement of acondition wherein shock waves defining the boundary between supersonicand subsonic conditions of gas flow are stabilized at a desired locationin an aerodynamic duct, given the mass flow, inlet Mach number, andpressure ratio for a gas being compressed. In general, variousstructures and/or systems as described herein may be used for startup—inorder to conduct the process of initiating operation and establishing astable shock system in aerodynamic ducts. In various embodiments,variable geometry inlets may be provided, allowing for a shock to beswallowed through a throat in an aerodynamic duct, to thereby start theaerodynamic duct. In other embodiments, aerodynamic ducts may beconfigured to allow external discharge of a portion of the gas flowthereto, in order to provide for startup, again by allowing a shock tobe swallowed through a throat in an aerodynamic duct. In otherembodiments, aerodynamic ducts may be configured to allow a portion ofthe gas flow thereto to internally bypass the throat. Such gas flow maybe reintroduced into a diverging portion of an aerodynamic duct. Thereduced gas flow through the throat of an aerodynamic duct allows forstarting of the aerodynamic duct. The performance of the aerodynamicducts when in a startup configuration would be roughly the same as mightbe found in an aerodynamic duct without adjustable gas flow and havingthe same effective contraction ratio (in other words, the degree ofblockage of the aerodynamic duct) for example, as in a fixed geometryaerodynamic duct. However, once startup is achieved and stablesupersonic flow is established, bypass valves, or gates, or otherstructures employed to provide for bypass of some gas around theconverging portion, or to provide reduced throat cross-sectional area,may be closed or returned to an operating position or operatingcondition. Thereafter, in an operational configuration, a compressor asdescribed herein provides aerodynamic ducts wherein a high pressureratio recovery is achieved even when a single stage of compression isemployed.

The term “un-start condition” may be used herein to describe a flowcondition under which gas to be compressed flows through an inlet muchless effectively than under compressor design conditions, and whereinsome, or even most of entering gas may be rejected from the inlet,instead of being properly ingested for effective operation of thecompressor. In various embodiments, during an un-start condition,supersonic flow conditions with stable shocks would not be properlyestablished at their design range locations within an aerodynamic duct.

The term “VGs” may be used to refer to vortex generators.

Turning now to FIG. 1, an exemplary design for a supersonic compressor40 is illustrated. The compressor 40 may utilize a rotor 42 having anaxis of rotation 44, and, for example a driving shaft 45, and aplurality of blades 46 extending into a gas flow passage 48. Blades 46may be sized and shaped to act on a selected incoming gas 50 to providea supersonic gas flow 52. A diffuser 54 is provided. In an embodiment,diffuser 54 may be disposed around a longitudinal axis 55 (shown withcenterline C_(L) in FIG. 1) and positioned to receive the supersonic gasflow 52. In an embodiment, the diffuser 54 may be provided as one ormore aerodynamic ducts 56. In some of the figures (see FIG. 15, forexample), the one or more aerodynamic ducts 56 may be individuallyfurther identified with a subscript as a first aerodynamic duct 56 ₁, asecond aerodynamic duct 56 ₂, a third aerodynamic duct 56 ₃, a fourthaerodynamic duct 56 ₄ (shown in FIG. 15), a fifth aerodynamic duct 56 ₅,and in FIG. 16, a sixth aerodynamic duct 56 ₆, and seventh aerodynamicduct 56 ₇, are shown for each individual aerodynamic duct 56 that may beutilized in a specific diffuser 54 design. More generally, a number N ofaerodynamic ducts 56 and a number B of blades 46 may be provided, withthe number B of blades 46 and the number N of aerodynamic ducts 56 beingunequal, in order to avoid adverse harmonic effects. While in variousprior art compressor designs a number B of blades 46 of N minus 1 (N−1)or of N plus one (N+1) has generally been considered acceptable to avoidadverse harmonic effects, it is noted herein that aerodynamic losses arereduced by minimizing the number of aerodynamic ducts 56, and morespecifically, by reducing the number of components exposed to asupersonic incoming gas flow stream. Thus, in an embodiment, the numberof blades 46 may considerably exceed the number of aerodynamic ducts 56,thereby reducing components exposed to supersonic flow. However anyratio between the number B of blades 46 and the number N of aerodynamicducts 56 should be selected to avoid adverse harmonic effects.

The aerodynamic ducts 56 each include a converging portion 58 and adiverging portion 60. In an embodiment, rotor 42 may be configured withblades 46 to turn incoming gas 50 to provide a supersonic relativevelocity gas flow 52 at a selected exit angle beta (β) relative to thecenterline C_(LD) of the one or more downstream aerodynamic ducts 56. Inan embodiment, but without limitation, the angle beta (β) may beprovided at zero degrees (0°), wherein the direction of gas flow 52 isaligned with the centerline C_(LD) of aerodynamic ducts 56, and thus aunique incidence angle is provided between the direction of the gas flow52 and the centerline C_(LD) of the one or more downstream aerodynamicducts 56. In other words, in an embodiment, a unique incidence angle isprovided since the direction of gas flow 52 matches the centerlineC_(LD) of an aerodynamic duct 56 into which the gas flow 52 occurs.However, it should be understood that configurations which are not soprecisely aligned may also be workable, but it must be noted that if theflow exit angle beta (β) is not aligned with respect to the aerodynamicducts 56, a series of shock waves or expansion fans (depending onwhether the relative angle of attack of the incoming flow is positive ornegative) will be formed to turn the flow to largely match the flowangle through the aerodynamic ducts 56 along centerline CLD. Such shockwave or expansion fan systems will result in total pressure loss whichwill contribute to a decrease in overall compression efficiency, andreduce the overall compressor ratio achieved for a given speed of blades46. As an example, a variation in the flow exit (or “incidence”) anglebeta (β) ranging from about 11.0 to about 8.0 degrees, at inflow Machnumbers of from about 2.0 to about 3.0, respectively, would result inabout three (3) percentage points of efficiency loss. Such increasedlosses and corresponding decreases in stage efficiency may be acceptablein various applications. However, in addition to shock wave or expansionfan conditions resulting in pressure and efficiency loss, adverse shockto boundary layer interaction, and or boundary layer separation issues,may arise as such off-design conditions become more severe, dependingupon the strength of the shock wave system and the thickness of theboundary layer system interacting therewith. And, adverse shock wave andaccompanying pressure signatures may be expected to reflect from blades46, especially at the trailing edges thereof, potentially increasingstress and reducing life of blades 46. Consequently, embodiments tendingto closely align flow exit angle beta (β) with the centerline C_(LD) ofaerodynamic ducts 56 should be considered optimal, although notlimiting.

A rotor 42 and a diffuser 54 together, as depicted in FIG. 1, provide astage of compression. In those cases where further compression isrequired, multiple stages of compression may be utilized in order toprovide gas at a desired final pressure, for example, as shown in FIG.34 below.

As shown in FIG. 1, in an embodiment, a diffuser 54 may include thereinone or more structures that enable startup of the shock wave, and one ormore structures that provide for control of boundary layer drag, as morefully addressed below. In an embodiment, bypass gas passageways 62 areprovided to remove a portion of incoming gas flow 52 during startupconditions, so as to adjust the effective contraction ratio of theassociated aerodynamic duct 56. In this manner, aerodynamic ducts 56 maybe designed for operation at high compression ratios, yet be adapted forstartup of a stable supersonic shock system within the aerodynamic duct56 that ultimately enables transition to high compression ratiooperation.

In an embodiment, aerodynamic ducts 56 may include one or more boundarylayer control structures, such as bleed ports 64 as seen in FIG. 1 forremoval of gas from aerodynamic ducts 56 as may be required for controlof boundary layer at surface 66 of the aerodynamic duct 56. As furtherdescribed below, boundary layer control may be provided by one or moreother or additional structures, such as via inlet jets 70 for energizinga boundary layer by gas injection (see FIGS. 26 and 27), and/or byvortex generators 72 or 74 (see, for example, FIGS. 20, 23, and 28).

Turning now to FIG. 2, as an example for a particular design and withoutlimitation, flow conditions are depicted for an embodiment for a designwithin a selected design envelope for a supersonic compressor. The rotor42 includes impulse blades 46, moving in the direction indicated byreference arrow 78. The use of impulse blades 46 in rotor 42 enablesefficient turning of the flow of an incoming gas, especially whenutilizing a rotor 42 having blades 46 with sharp leading edges 80 andsharp trailing edges 94. At location A, upstream of rotor 42, a smalltangential velocity (as compared to tangential velocity after exit fromrotor 42 as described below) may be encountered prior to the rotor 42,as indicated by the velocity diagram shown in cloud 82. At the entryplane to the rotor 42, i.e. at location B, the gas velocity isaccelerated as indicated in the velocity diagram shown in cloud 84. Atthe exit plane the rotor 42, i.e. at location C, the gas has beenpartially accelerated and is moving as indicated in the velocity diagramshown in cloud 86. Finally, after exit from rotor 42, at location D, thegas velocity is as indicated in the velocity diagram shown in cloud 88.Basically, an impulse bladed rotor 42 allows a high degree of turning ofthe incoming gas 50, through an angle alpha (α). Moreover, as seen inthe velocity vector diagram set forth in cloud 88, the vector sum of theaxial velocity at location D (V_(D) _(—) Axial of about 628 feet persecond), the tangential velocity at location D (V_(D) _(—) Tangential ofabout 2004.3 feet per second), provides an overall relative velocity ofgas stream 52 at location D (V_(D) of about 2157.6 feet per second),which is thus supersonic for gas stream 52 as it enters an aerodynamicduct 56 of diffuser 54. Thus, as seen in FIG. 1, in an embodiment, thedesired supersonic velocity of gas stream 52 entering the aerodynamicducts 56 of diffuser 54 is achieved by a combination of velocity of gasthrough the blades 46 and the tangential rotation of the rotor 42.

In an embodiment for a supersonic compressor 40 such as illustrated inFIG. 1, the selected inlet gas passing through the blades 46 of therotor 42 may be turned by an angle alpha (α) of at least ninety (90)degrees. In an embodiment of compressor 40, the selected inlet gaspassing through the rotor 42 may be turned by an angle alpha (α) of atleast one hundred (100) degrees. In an embodiment of compressor 40, theselected inlet gas passing through the blades 46 of the rotor 42 may beturned by an angle alpha (α) of at least one hundred ten (110) degrees.In an embodiment of compressor 40, the selected inlet gas passingthrough the blades 46 of the rotor 42 may be turned by an angle alpha(α). The angle alpha (α) may be at least ninety (90) degrees, forexample, between about ninety (90) degrees and about one hundred twentyfive (125) degrees, or between about ninety (90) degrees and about onehundred sixty (160) degrees, or between about one hundred twelve (112)degrees and about one hundred fourteen (114) degrees. Details ofexemplary designs for various impulse type blades for use in supersoniccompressors may be found by those of skill in the art from varioussources. One helpful reference may include a NASA report entitled“Analytical Investigation of Supersonic Turbomachinery Blading—SectionII—Analysis of Impulse Turbine Blade Sections”, by Louis J. Goldman,Published as Report No. NASA-TN-D-4422, on Apr. 1, 1968, which isincorporated herein by reference, and to which the reader may refer foradditional background in implementing an impulse blade in a supersoniccompressor design as further taught herein.

As shown in FIG. 3, in an embodiment, each of the plurality blades 46 inrotor 42 may have a hub end 90, a tip end 92, and a trailing edge 94. Inan embodiment, the blades 46 are provided with supersonic gas flow 52 attheir trailing edge 94. In an embodiment, supersonic gas flow at thetrailing edge 94 may be from the hub end 90 to the tip end 92 of thetrailing edge 94.

As shown in FIGS. 10 and 11, in an embodiment, a rotor 100 may beprovided having a shroud 102 for blades 103. Such shrouded blades 103 onrotor 100 will be understood by those of skill in the art to beotherwise as just noted above as regards supersonic gas flow on blades103 from a hub end 104 to a tip end 106 at trailing edge 108. In anembodiment, shroud 102 may include labyrinth seal portions 110 and 112.By use of a labyrinth seal or other suitable seal, such as a honeycombseal, a dry gas seal, brush seals, etc, the rotor 100 may be effectivelysealed with respect to a downstream aerodynamic duct such as duct 162,so as to minimize gas leakage during flow therebetween.

A cross-sectional view of a diffuser 54 is shown in FIG. 5, taken acrosssection line 5-5 in FIG. 1. As shown in that embodiment, five (5)aerodynamic ducts are utilized, more specifically aerodynamic ducts 56₁, 56 ₂, 56 ₃, 56 ₄, and 56 ₅, each having a converging portion 58 and adiverging portion 60. Inlet bypass gas passageways 62 are shown, asuseful for starting, by removal of discharge gas 113 as functionallyillustrated in FIG. 6 and as further discussed below with reference toFIG. 6. As illustrated in FIGS. 5 and 6, sub-chambers 114 ₁, 114 ₂, 114₃, 114 ₄, and 114 ₅ are shown as transport conduits for discharge gas113 from respective associated aerodynamic ducts 56 ₁, 56 ₂, 56 ₃, 56 ₄,and 56 ₅. Boundary layer control structures, here provided in the formof boundary layer bleed ports 64, are shown for use in boundary layercontrol, by removal of bleed gas 121. Boundary layer bleed sub-chambers122 ₁, 122 ₂, 122 ₃, 122 ₄, and 122 ₅ are shown as transport conduitsfor boundary layer bleed gas 121 from respective associated aerodynamicducts 56 ₁, 56 ₂, 56 ₃, 56 ₄, and 56 ₅. In general, the sub-chambers 114₁, 114 ₂, 114 ₃, 114 ₄, and 114 ₅ for handling discharge gas 113 andsub-chambers 122 ₁, 122 ₂, 122 ₃, 122 ₄, and 122 ₅ for handling bleedgas 121 may be located inwardly from their respective aerodynamic ducts56 from which such discharge gas 113 and bleed gas 121 are removed.

Turning now to FIG. 6, an enlarged detail of a portion of an exemplaryaerodynamic duct 56 ₁ is illustrated. Here, the use of inlet bypass gaspassageways 62 is shown, as useful for removal of discharge gas 113during starting of the compressor. Also, boundary layer bleed ports 64are shown for use in boundary layer control by removal of bleed gas 121to a sub-chamber 122 ₁, which control may occur during normal operation,or during starting, or both. Further, an exemplary vortex generator 74is shown within the aerodynamic duct 56 ₁ for control of a boundarylayer by mixing the boundary layer flow with higher velocity gas flowmore distant from surface 123. Multiple vortex generators, whether ofthe specific designs described herein, or chosen from one or more vortexgenerator configurations heretofore known to those of skill in the art,may be utilized as appropriate in any particular design.

In an embodiment, such as illustrated in FIG. 4, bypass gas passageways62 may include, in fluid communication therewith, outlet valving 116positionable between an open, startup condition wherein discharge gas113 is passed therethrough, and a closed, operating condition whichminimizes or stops flow of discharged bypass gas 113. A sub-chamber 114may be provided for collection of bypass gas 113, with the outletvalving 116 regulating passage of such collected bypass gas 118 outwardvia external passageways 120. In such embodiment, the aerodynamic ducts56 have outlets in the form of bypass gas passageways 62 that arefluidly connected to external passageways 120. In an embodiment,collected bypass gas 118 may be returned as shown by broken line 118′ toinlet passageway 48. Or, in the case of compression of air, collectedbypass gas 118 may be discharged directly to the atmosphere, asindicated by broken line 119 in FIGS. 4 and 14.

Similarly, in various embodiments, the boundary layer bleed ports 64 mayinclude outlet valving 124 positionable between an open position whereinbleed gas 121 is passed therethrough (see FIG. 4), and a closed positionwhich avoids boundary layer gas removal via removal of bleed gas 121.For example, a boundary layer bleed sub-chamber 122 is shown forcollection of bleed gas 121, with outlet valving 124 for passage ofcollected bleed gas 126 outward by external line 128. In suchembodiment, the boundary layer bleed ports 64 from aerodynamic ducts 56are fluidly connected to external lines 128. As also shown in FIG. 4, inan embodiment, collected bleed gas 126 may be recycled, optionally shownby broken line 126′, and returned to inlet passageway 48. Or, in case ofcompression of air, the collected bleed gas 126 may be discharged to theatmosphere, as indicated by broken line 127 in FIGS. 4 and 14.

In other embodiments, as seen in FIGS. 13 and 14, a compressor may beprovided using internal starting bypass gas passageways 130 as definedby internal walls 131 of an internal gas passageway housing 133. In suchconfiguration, the internal bypass gas passageways 130 are fluidlyconnected internally within or adjacent the aerodynamic ducts 132, toallow bypass gas 134 to escape a converging portion 136, and return thebypass gas 134 directly to the aerodynamic duct 132, as shown byreference arrow 148 in FIG. 13, to the diverging portion 138 thereof. Inan embodiment, a hinged inlet door 140 may be provided with actuatorlinkage 142 for opening a bypass outlet 144 shown in broken lines.Bypass gas 134 escapes through bypass outlet 144 and is then returned asindicated by reference arrows 146 and 148 in FIG. 13 through bypassreturn opening 154. A hinged return door 150 may be provided withactuator linkage 152 for opening a bypass return opening 154 shown inbroken lines in FIG. 13.

Attention is directed to FIG. 13A, which shows yet another embodimentfor achieving startup of a supersonic shock wave in an aerodynamic duct132. In FIG. 13A, a bypass outlet door 155 provides a bypass outletopening 156 shown in broken lines between end walls 156 ₁ and 156 ₂ toallow gas shown by reference arrows 157 to escape the converging portion136 of the aerodynamic duct 132. In an embodiment, an actuator 158 maybe provided to move back and forth as noted by reference arrows 158 ₁(to open), and 158 ₂ (to close) bypass outlet door 155, using linkage158 ₃ to pivot bypass outlet door 155 about pivot pin 155 ₁. Escapingbypass gas noted by reference arrow 157 ₁ is contained by bypass gaspassageway wall 159, which provides a pressurizable plenum to containbypass gas. In an embodiment, actuator 158 is not a bounding wall, asthe escaping bypass gas noted by reference arrow 157 ₁ is free to passas indicated by reference arrows 161 _(A) and 161 _(B) outward to bypassgas passageway wall 159. Once pressurized, the bypass gas then escapesthrough the enlarged throat opening O₂, and thence downstream of thethroat opening O₂ as indicated by reference arrow 157 ₂. An enlargedarea of A₂ (not shown but corresponding to opening at throat O₂) ofthroat O₂ when in a startup configuration (as compared to an area of A₁of throat O₁ when in an operational configuration) enables downstreampassage through the aerodynamic duct 132 of bypass gas as indicated byreference arrow 157 ₂. In an embodiment, the bypass outlet door 155 maybe provided with boundary layer bleed passages 155 ₂, for boundary layerbleed as noted by reference arrows 155 ₃ and 155 ₄. More generally,startup of the supersonic shock wave is established by opening up bypassgas passageways such as bypass outlet door 155, and then bringing theblades 46 up to full speed. Then, the bypass outlet door may be smoothlyclosed to bring the throat O₁ of aerodynamic duct 132 into a design areacondition which establishes a design contraction ratio for aerodynamicduct 132. At that point, back pressure, that is the static pressure indiverging portion 138 of aerodynamic duct 132, is allowed to rise toestablish the design discharge pressure for operation. Boundary layercontrol structures are utilized during operation to control boundarylayers, whether by bleed, mixing, injection, combinations thereof, orother suitable means. For shutdown, back pressure is reduced, and drivefor blades 46 is turned off, and the compressor is allowed to spin to astop.

Turning to FIGS. 11 and 12, a compressor may be provided in anembodiment using geometrically adjustable portion(s) 160 in anaerodynamic duct 162. As seen in FIGS. 11 and 12, a geometricallyadjustable portion 160 may be positionable between a location for use ina startup condition shown in broken lines, with a larger throat O₂ areaof A₂, wherein the converging portion 164 allows increased flow of aselected gas through the aerodynamic duct 162, and a location for use inan operating condition in which the converging portion 164 is set to aselected operating position, shown in solid lines, with a throat O₁ areaof A₁. The adjustment of geometrically adjustable portion 160 to theoperating position and thus providing a smaller throat O₁ area A₁ shownin solid lines in FIG. 12 allows operation with a higher compressionratio than when geometrically adjustable portion 160 is at the startupposition indicated in FIG. 12 by broken lines 163 and providing throatO₂ area A₂. In other words, the geometrically adjustable portion(s) 160move, to change the contraction ratio of an aerodynamic duct 162. Invarious embodiments, one or more geometrically adjustable portions 160may be located in one or more of aerodynamic ducts 162, as provided fora particular compressor. As indicated in FIG. 12, in an embodiment,adjustment of a geometrically adjustable portion 160 may includeextending the length of the converging portion 164 and diverging portion165 of a duct 162 by a length L. In an embodiment, such adjustment maybe achieved by use of a pivot pin 167. In an embodiment, an actuator166, extending between an anchor 168 and an attachment point 170, may beprovided to move the geometrically adjustable portion 160 and allowmovement such as at pivot pin 167.

Returning now to boundary layer control structures, in an embodiment,such structures may be configured as boundary layer bleed ports 64 inthe various aerodynamic ducts 56 in diffuser 54, such as shown in FIG.1, or in FIG. 4. Such boundary layer bleed ports 64 may be provided byperforations in one or more bounding walls, such as in surface 66 of adiverging portion 60 in an aerodynamic duct as shown in FIG. 1 or 4.Adjacent the boundary layer bleed ports 64 may be bleed sub-chambers,such as sub-chamber 122 noted above with respect to the embodimentdepicted in FIG. 4, or as may be seen in FIG. 13. Thus, a bleedsub-chamber 122 may be provided in fluid communication with boundarylayer bleed ports 64, and thus bleed sub-chambers 122 are configured forpassage therethrough of gas removed through the boundary layer bleedports 64. Although the boundary layer bleed ports 64 are shown in adiverging portion 60, such bleed ports may be located in other boundingwalls of aerodynamic ducts 56, such as on radially outward portions, oron sidewalls, or on other radial inward portions.

In yet another embodiment, boundary layer control may be provided viause of boundary layer control structures such as inlet jets 70 as shownin FIGS. 26 and 27. (Note that inlet jets 70 may also be described asinlet nozzles.) In an embodiment inlet jets 70 may be oriented to injectgas 172 into a boundary layer 174 in a direction consistent with flow ofgas through one or more aerodynamic ducts 56, thus speeding up andthereby energizing the boundary layer 174 in the gas flow direction,which is shown by reference arrow 176 in FIGS. 26 and 27. As shown inFIG. 26, in an embodiment, injection gas chambers 180 defined by chamberwalls 182 may be provided adjacent the one or more aerodynamic ducts 56.The injection gas chambers 180 are in fluid communication with inletjets 70, and injection gas chambers 180 are configured for passagetherethrough of gas to be injected via the inlet jets 70. Thus in anembodiment, the boundary layer control structures configured as inletjets 70 are positioned adjacent to a bounding surface 184 or 185 in oneor more aerodynamic ducts 56. As shown in FIGS. 26 and 27, the inletjets 70 may be positioned to discharge gas 172 in a directionsubstantially aligned with flow of gas 176 through the one or moreaerodynamic ducts 56. As depicted in FIG. 27, the injection inlet jets70 are sized and shaped to provide a jet of gas 172 that energizes theboundary layer by increasing the momentum of an adjacent flow ofboundary layer 174 of gas of the aerodynamic ducts 56 into which gas 172from the injection inlet jet 70 is injected.

In an embodiment, injection jet(s) 70 may be provided in the form of atleast one nozzle in fluid communication with a source of high pressuregas, such as injection gas chambers 180. Gas from the source of highpressure gas such as injection gas chambers 180 is provided at apressure higher than the pressure of the gas in the boundary layer 174.The injection jets 70 have an outlet nozzle 183 downstream of surface184 and adjacent a surface 185 in an aerodynamic duct. The injectionjets 70 are positioned and shaped in a manner so as to direct the highpressure gas from the source of high pressure gas out through theinjection jets 70 and into the boundary layer 174. In an embodiment, theinjection jets 70 may be shaped in a manner so as to direct such highpressure gas both into the boundary layer 174 and along the surface 185,to re-energize the boundary layer's 174 pressure profile, so that suchpressure profile approaches a freestream gas profile prior to ingestionof the boundary layer gas. In an embodiment, the surface 185 may besubstantially smooth and continuous surface downstream of the injectionjets 70.

Turning now to FIGS. 20 through 25, in an embodiment, boundary layercontrol structures may be provided as vortex generators, such as vortexgenerators 72 and 74. Further, as shown in FIG. 11, a vortex generator72 may be located on a converging portion 164 in an aerodynamic duct162. Likewise, a vortex generator 74 may be located on a divergingportion 165 of an aerodynamic duct 162. As shown in FIG. 20, the vortexgenerator 72 may include a base 200 attached to a suitable surface 201with a forward end 202 and a leading edge 204 extending outward andrearward. i.e., in a downstream direction from the forward end 202 ofthe base to an outward end 206. In an embodiment, the leading edge 204includes at least one angular discontinuity 210 along the leading edge204, for generating at least one vortex. In an embodiment, the leadingedge 204 includes a first angular discontinuity 210 at a height H₁ abovethe base 200, and a second angular discontinuity 212 at a height H₂above the base 200, for generating two vortices. As shown for vortexgenerator 74 in FIG. 23, in an embodiment, the leading edge 204 includesa first angular discontinuity 210 at a height H₁ above the base 200, asecond angular discontinuity 212 at a height H₂ above the base 200, anda third angular discontinuity 214 at a height H₃ above the base 200, forgenerating three vortices. In various embodiments, a plurality of vortexgenerators 72 and or 74 may be provided in each of one or moreaerodynamic ducts 162 (see FIG. 12), or like aerodynamic ducts 56 asillustrated, for example, in FIG. 1. Vortex generators may be providedin the just described novel configurations, or in heretofore knownconfigurations as will be understood by those of skill in the art, or infurther variations of either.

In an embodiment, vortex generators may be provided having height H₁that is about 1.6 times the result of height H₂ minus height H₁. In anembodiment, height H₂ may be about 1.6 times the result of height H₃minus height H₂. Thus, in an embodiment, the height ratios ofdiscontinuities in vortex generators for generating vortices in therespective multi-vortex embodiments may be about 1.6, roughly the socalled “golden ratio”. Generally, the golden ratio (more precisely1.618) is denoted by the Greek lowercase letter phi (φ). With respect tovortex strength, if the height ratios are equal to phi (φ), then thestrength ratios, that is the comparative strength between the first andsecond vortices, should be equal to (φ)⁻². Generally, as depictedbetween FIGS. 21 and 22, and likewise in FIGS. 24 and 25, in a vortexgenerator design, a useful technique may be to use the larger, andstronger vortex, say V₁, to turn a smaller vortex, say, V₂, toward thesurface 201. Likewise, with three vortices, such technique involvesturning the larger and stronger vortices, say V₁ and V₂, to drive thesmaller vortex V₃ toward the surface 201. In such manner, a largervortex V₁, which might not otherwise be able to mix with a boundarylayer against surface 201, is able to bring energy to mix higher energyfluid with the boundary layer by virtue of carriage of the smallervortex V₃ toward surface 201.

Further variations in suitable vortex generator designs are illustratedin FIGS. 35 through 40. In various vortex generator configurations, oneor more vortex generators may be provided, such as the two vortexgenerators 190 and 191 that are shown in FIG. 35. The vortex generators190 and 191 each generate a single vortex, V₁₉₀ and V₁₉₁, respectively.As shown in FIG. 35, vortex generators 190 and 191 are located in thediverging portion 60 of aerodynamic duct 56. Alternately as suggested inFIG. 37 (or additionally), vortex generators 190 and 191 may be locatedin a converging portion 58 of an aerodynamic duct 56.

As noted in FIG. 36, each of the one or more vortex generators such asvortex generators 190 and 191 have a base 193 with a forward end 194 anda leading edge 195 extending outward to an outward end 196. In anembodiment, a leading edge 195 may comprise an angular discontinuity toenable the vortex generators 190 or 191 to each generate a singlevortex, such as the vortices V₁₉₀ or V₁₉₁ as shown in FIG. 35, whichvortices are configured to control a boundary layer (not shown) of thegas flow 52 in aerodynamic duct 56. As may be noted in FIG. 36, in anembodiment, the outward end 196 may be located at a height H_(Z) abovethe base 193. In an embodiment such as shown in FIG. 36, an aerodynamicduct 56 has a height H_(D), and the height H_(Z) of the outward end 196above the base 193 is about fifty percent (50%), or less, of the heightH_(D). For example, and without limitation, in an embodiment, the heightH_(D) of aerodynamic duct 56 may be configured to be about 0.264 inches,and the height H_(Z) of the outward end 196 above base 193 may be about0.133 inches.

Additional details of a workable configuration for vortex generators 190and 191 are provided in FIG. 37. It can be seen that the vortexgenerators 190 and 191 may be offset by an angle sigma sub one (Σ₁) andsigma sub two (Σ₂), respectively from a longitudinal centerline C_(LD)of aerodynamic duct 56. In an embodiment, angle sigma sub one (Σ₁) andsigma sub two (Σ₂) may each be about twenty degrees (20°). In anembodiment, a gap of width W_(G) may be provided between the upstream,forward ends 194 of vortex generators 190 and 191, of about zero pointthree three (0.33) inches. In an embodiment the vortex generators 190and 191 may be provided in a suitable length to achieve the sizing ofvortices V₁₉₀ and V₁₉₁ that may achieve the desired boundary layercontrol results. In an embodiment, and without limitation, vortexgenerators 190 and 191 may be provided in a length L₁₉₀, for example ofabout 0.53 inches. In such an embodiment, and without limitation, vortexgenerators 190 and 191 may be provided having a width W₁₉₀ of base 193of about 0.07 inches.

Turning now to FIG. 38, an embodiment is illustrated wherein anaerodynamic duct 56 is provided with four vortex generators, 186, 187,188, and 189. Each of vortex generators 186, 187, 188, and 189 generatea single vortex, namely V₁₈₆, V₁₈₇, V₁₈₈, and V₁₈₉, respectively.

As noted in FIG. 39, each of the one or more vortex generators such asvortex generators 186, 187, 188, and 189 have a base 193 with a forwardend 194 and a leading edge 195 extending outward to an outward end 196.In an embodiment, a leading edge 195 may comprise an angulardiscontinuity to enable the vortex generators 186, 187, 188, and 189 toeach generate a single vortex, such as the vortices shown in FIG. 38,which vortices are configured to control a boundary layer (not shown) ofthe gas flow 52 in aerodynamic duct 56. As may be noted in FIG. 39, inan embodiment, the outward end 196 may be located at a height H_(Z)above the base 193. In an embodiment such as shown in FIG. 36, wherefour vortex generators 186, 187, 188, and 189 are used, an aerodynamicduct 56 may be provided having a height H_(D), and the height H_(Z) ofthe outward end 196 of the vortex generators 186, 187, 188, and 189above the base 193 is about twenty five percent (25%), or less, of theheight H_(D). For example, and without limitation, in an embodiment, theheight H_(D) of aerodynamic duct 56 may be configured to be about 0.264inches, and the height H_(Z) of the outward end 196 above base 193 maybe about 0.067 inches.

Additional details of a workable configuration for vortex generators186, 187, 188, 189 are provided in FIG. 40. It can be seen that thevortex generators 186 and 187 may be offset by an angle sigma sub three(Σ₃) and sigma sub four (Σ₄), respectively from a longitudinalcenterline C_(LD) of aerodynamic duct 56. In an embodiment, angle sigmasub three (Σ₃) and sigma sub four (Σ₄) may each be from about fifteendegrees (15°) to about twenty degrees (20°). In an embodiment, a gap ofwidth W_(G) may be provided between the upstream, forward ends 194 ofvortex generators 186 and 187, of about zero point six six (0.66)inches. In an embodiment the vortex generators 186, 187, 188, and 189may be provided in a suitable length to achieve the sizing of vorticesV₁₈₆, V₁₈₇, V₁₈₈ and V₁₈₉ that may achieve the desired boundary layercontrol results. In an embodiment, and without limitation, vortexgenerators 186, 187, 188, and 189 may be provided in a length L₁₈₇, forexample of about zero point two six six (0.266) inches. In such anembodiment, and without limitation, vortex generators 186, 187, 188, and189 may be provided having a width W₁₈₇ of base 193 of about 0.07inches.

In various embodiments, as shown in FIGS. 17, 18 and 19, the one or moreaerodynamic ducts 56 are disposed in a stator, such as a stationarydiffuser 54 as depicted in FIG. 1 above, and may be wrapped around alongitudinal axis, shown along the centerline C_(LS). In an embodiment,as indicated in FIG. 17, one or more of the one or more aerodynamicducts 56 of a stationary diffuser 221 are wrapped as if over asubstantially cylindrical substrate 220. In such an embodiment,aerodynamic ducts 56 may be helically arranged in adjacent positions ata substantially constant helical angle psi (Ψ) about the longitudinalaxis shown along the centerline of the stator, C_(LS). Alternately, theorientation of aerodynamic ducts 56 may be described by use of thecomplementary lead angle delta (Δ), as shown in FIG. 17. In such anembodiment, the centerline C_(LD) of a first aerodynamic duct 56 ₁ andthe centerline C_(LD) of a second aerodynamic duct 56 ₂ (and other ductsin the embodiment) may be parallel. In various embodiments, a helicalangle psi (Ψ) in the range of from about forty-five degrees (45°) toabout eighty degrees (80°) may be employed. In the designs disclosedherein, it may be advantageous to receive gas in aerodynamic ducts, forexample, 56 ₁ in FIG. 17, without turning the flow as delivered fromblades 46 as shown in FIG. 1. In a different design as depicted in FIG.18, aerodynamic ducts 56 ₃ and 56 ₄ of a diffuser 223 may be wrapped asif over an outwardly expanding conical section as a substrate 222. Inyet another and still different alternative embodiment, as seen in FIG.19, aerodynamic ducts 56 ₆ and 56 ₇ in a diffuser 225 may be wrapped asif over an inwardly decreasing conical section as a substrate 224.

Overall, as may be envisioned in part from FIG. 9, a supersonic gascompressor 230 may be provided for compressing a selected gas 232, wherethe compressor 230 includes a casing 234 having a low pressure gas inlet236 and a high pressure gas exit 238. A volute or collector 239 may beutilized downstream of the diffuser 54 to further convert kinetic energyto pressure energy in a high pressure gas 240. A rotor 42 with blades 46as shown in FIG. 1 (or shrouded blades 103 on rotor 100 as shown inFIGS. 10 and 11) may be provided to act on the selected gas 232 toimpart velocity thereto to provide a supersonic gas flow 52 (see FIG. 1)to a diffuser 54 that includes one or more aerodynamic ducts 56. Asshown in FIG. 9, provision may also be made for a deswirler 57, locateddownstream of diffuser 54, to turn the gas flow toward the axialdirection, when required. However, losses associated with deswirler 57may be avoided in some instances, and may be optional when dischargingto a volute 239 as indicated in FIG. 9. The rotor 42 with blades 46 (orshrouded blades 103 on rotor 100, as shown in FIGS. 10 and 11) may bedriven by shaft 238 from driver 241 (e.g., electric motor or other powersource), the choice of driver type and size, and associated drive traincomponents such as gearbox 242 or bearings 244, etc., may be selected bythose of skill in the art for a particular application.

As can be seen in FIG. 28, an exemplary aerodynamic duct 56 ₈ may beprovided in a stationary diffuser 54 of the type shown in FIG. 1. Theaerodynamic duct 56 ₈ shown in FIG. 28 may be considered, in anembodiment, as generally helically disposed about a longitudinal axis,such as about centerline C_(LS) of FIG. 17. Returning to FIG. 28, theaerodynamic duct 56 ₈ includes a converging portion 58 and a divergingportion 60 (here shown provided by ramps 246 and 248, respectively, onthe radially-inward side of the aerodynamic duct 56 ₈) that with inputof a supersonic (Mach>1) gas flow generates a plurality of oblique shockwaves S₁ to S_(x) and a normal shock wave S_(N) in a selected gas 50 asthe gas passes through the aerodynamic duct 56 ₈ from supersonicconditions (Mach>1) to subsonic conditions (Mach<1). The aerodynamicduct 56 ₈ may be designed, i.e., sized and shaped, for an inlet relativeMach number for operation associated with a design operating pointselected within a design operating envelope for a selected gascomposition, gas quantity, and gas compression ratio. A compressordesign may be configured for a selected mass flow, that is for aparticular quantity of gas that is to be compressed, and that gas mayhave certain inlet conditions with respect to temperature and pressure(or an anticipated range of such conditions), that must be considered inthe design. The incoming gas may be relatively pure single component, ormay be a mixture of various elements or various compounds or of variouselements and compounds, or the gas may be expected to range incomposition. And, it may be desired to achieve a particular finalpressure, when starting at a given inlet gas pressure, and thus, adesired gas compression ratio must be selected for a particularcompressor design. Given design constraints such as gas composition,mass flow of gas, inlet conditions, and desired outlet conditions theaerodynamic ducts for a particular compressor must be sized and shapedfor operation at a selected inlet Mach number and gas compression ratio.The designs described herein allow use of high gas compression ratios,especially compared to self starting compressor designs that lack theability to adjust the effective contraction ratio. Thus, the designsprovided herein provide for compression in aerodynamic ducts which canbe started, as regards swallowing a shock structure and establishing astable supersonic shock configuration during operation, yet retaindesign features that enable high pressure ratio operation, includingoblique shock structure and throat size to support design throughput andcompression pressure ratios. As shown in FIG. 28, in an embodiment,bypass gas passageways may be provided as outlets 250 in a boundingsurface 252 of aerodynamic duct 56 ₈ (here bounding surface 252 is shownas a radially outward bounding surface of aerodynamic duct 56 ₈). Thebypass gas passageway outlets 250 are in fluid communication withoutboard chambers 254 (individually indicated as outboard chambers 254₁, 254 ₂, 254 ₃, and 254 ₄) so that the effective contraction ratio ofaerodynamic duct such as duct 56 ₈ may be changed by removal of gastherefrom, as indicated by arrows 256. Also, in an embodiment, asuitable boundary layer control structure as described herein may beselected, such as the use of a plurality of vortex generators 72, 74.

Similarly, as can be seen in FIG. 29, an aerodynamic duct 56 such asexemplary aerodynamic duct 56 ₉ may be provided in a stationary diffuser54 such as shown in FIG. 1. The aerodynamic duct 56 ₉ shown in FIG. 29may be helically wrapped around a longitudinal axis of a diffuser 54,for example as if the section provided in the present FIG. 29 were takenalong the centerline C_(LD) shown for aerodynamic duct 56 ₂ in FIG. 1above. As seen in FIG. 29, the aerodynamic duct 56 ₉ includes aconverging portion 58 and a diverging portion 60 (here shown provided byramps 260 and 262, respectively, on the radially outward side ofaerodynamic duct 56 ₉) that with input of a supersonic (Mach>1) gas flowgenerates a plurality of oblique shock waves S₁ to S_(x) and a normalshock wave S_(N) in a selected gas 50 as the gas passes through theaerodynamic duct 56 ₉ from supersonic conditions (Mach>1) to subsonicconditions (Mach<1). The aerodynamic duct 56 ₉ may be designed, i.e.,sized and shaped, for an inlet relative Mach number for operationassociated with a design operating point selected within a designoperating envelope for a selected gas composition, gas quantity, and gascompression ratio. As shown in FIG. 29, in an embodiment, bypass gaspassageways may be provided as outlets 264 in a bounding surface 266 ofaerodynamic duct 56 ₉ (here bounding surface 266 is shown as a radiallyinward bounding surface of aerodynamic duct 56 ₉). The bypass gaspassageway outlets 264 are in fluid communication with inboardsub-chambers 268 (individually indicated as inboard sub-chambers 268 ₁,268 ₂, 268 ₃, 268 ₄, etc.) so that the effective contraction ratio ofaerodynamic ducts such as duct 56 ₉ may be changed by removal of gastherefrom, as indicated by arrows 269. Also, in an embodiment, asuitable boundary layer control structure as described herein may beselected, for example, using boundary layer bleed ports 270 for removalof gas 271 into inboard bleed sub-chambers 272 ₁, 272 ₂, 272 ₃. Also, aplurality of vortex generators 72, 74, may, in an embodiment, beemployed for assistance in boundary layer control. However, note theavailability of outboard chambers 273 ₁, 273 ₂, 273 ₃, etc., which alsomay be utilized as otherwise described herein for either bypass gasremoval, or for boundary layer bleed and control, as appropriate for aparticular design.

Yet another configuration for an exemplary aerodynamic duct 56 ₁₁ foruse in a diffuser 54 such as first shown in FIG. 1 may be seen in FIG.30. As shown in FIG. 30, an exemplary aerodynamic duct 56 ₁₁ may, in anembodiment, be in a helical arrangement and wrapped around alongitudinal axis of a diffuser, for example as if the section providedin the present FIG. 30 were taken along the centerline C_(LD) shown foraerodynamic duct 56 ₂ in FIG. 1. As shown in FIG. 30, the aerodynamicduct 56 ₁₁ includes a converging portion 58 and a diverging portion 60.In this embodiment, opposing radial bounding walls in the form ofinboard converging ramp 274 and outboard converging ramp 276 provide aconverging portion 58. Also, in this embodiment, opposing radialbounding walls in the form of an inboard diverging ramp 280 and anoutboard diverging ramp 281 provide a diverging portion 60.Additionally, the aerodynamic duct 56 ₁₁, like other aerodynamic ducts56, includes sidewalls as necessary to form a pressurizable duct, whichin an embodiment may be in the form of lateral partition walls, notshown in FIG. 28, 29, or 30, but may be provided as partition walls 364(individually identified as partition walls 364 ₁, 364 ₂, 364 ₃, etc. asappropriate given the number of aerodynamic ducts 56 utilized) asillustrated for exemplary diffuser 54 designs shown in FIGS. 1, 7, and8. In the embodiment shown in FIG. 30, the inboard converging ramp 274and the outboard converging ramp 276 receive input of a supersonic(Mach>1) gas flow and generate a plurality of oblique shock waves S₁ toS_(x) and a normal shock wave S_(N) in a selected gas 50 as the gaspasses through the aerodynamic duct 56 ₁₁ from supersonic conditions(Mach>1) to subsonic conditions (Mach<1). The aerodynamic duct 56 ₁₁ maybe designed, i.e., sized and shaped, for an inlet relative Mach numberfor operation associated with a design operating point selected within adesign operating envelope for a selected gas composition, gas quantity,and gas compression ratio. As shown in FIG. 30, in an embodiment, bypassgas passageways may be provided as outlets 278 in a bounding surface ofaerodynamic duct 56 ₁₁ (here a radially outward bounding surface ofaerodynamic duct 56 ₁₁ is shown as outboard converging ramp 276). Thebypass gas passageway outlets 278 are in fluid communication withoutboard chambers 282 (individually indicated outboard chambers 282 ₁,282 ₂, 282 ₃, etc.) so that the effective contraction ratio ofaerodynamic ducts such as duct 56 ₁₁ may be changed by removal of gastherefrom, as indicated by arrows 284. Additionally, bypass gaspassageways may be provided as outlets 288 in a bounding surface ofaerodynamic duct 56 ₁₁ (here such bounding surface is shown as aradially inward bounding surface of aerodynamic duct 56 ₁₁, namelyinboard converging ramp 274). The bypass gas passageway outlets 288 arein fluid communication with inboard sub-chambers 292 (individuallyindicated as inboard sub-chambers 292 ₁, 292 ₂, and 292 ₃, etc.) so thatthe effective contraction ratio of aerodynamic ducts such as duct 56 ₁₁may be changed by removal of gas therefrom, as indicated by arrows 284and 294 (shown in FIG. 30). Also, in an embodiment, a suitable boundarylayer control structure as described herein may be selected, forexample, using boundary layer bleed ports 294 for removal of gas 296into inboard bleed sub-chambers 298 ₁, 298 ₂, 298 ₃. And, in anembodiment, a plurality of vortex generators, indicated as “VGs” 72, 74,may be utilized to minimize adverse boundary layer effects.

Attention is now directed to FIGS. 31, 32, and 33, which provide yetfurther embodiments for a supersonic compressor, and more specifically,for the configuration of a stationary diffuser in such a compressor,wherein lateral gas compression, which occurs in a channel betweenbounding adjacent sidewalls, rather than radial compression (i.e., whichoccurs in a channel between radially spaced apart bounding walls), isutilized in an aerodynamic duct. FIGS. 31, 32, and 33 provide partialcircumferential views showing the longitudinal centerline C_(LS) of astationary diffuser (stator), and generally helical aerodynamic ductsused therein, as well as the accompanying rotor 42 (as is seen inFIG. 1) and its rotational centerline 299. In the embodiments depictedin FIGS. 31, 32, and 33, an aerodynamic duct design is provided whereinthe compression is done laterally, that is, in a channel between spacedapart sidewalls, rather than via radially spaced apart bounding wallsthat occur in an aerodynamic duct, as for example are shown in FIG. 1,or as just set forth in detail in FIGS. 28, 29, and 30 for variousalternate embodiments. FIG. 31 illustrates an embodiment whereincompression is performed in aerodynamic ducts 300 (individuallyindicated as aerodynamic ducts 300 ₁, 300 ₂, and 300 ₃, etc.) using arespective downstream sidewall 302. FIG. 32 provides an embodimentwherein compression is performed in aerodynamic ducts 304 ₁, 304 ₂, and304 ₃, etc using an upstream sidewall 306. FIG. 33 provides anembodiment wherein compression is performed in aerodynamic ducts 308 ₁,308 ₂, and 308 ₃, etc using both a downstream sidewall 310 and anupstream sidewall 312.

In FIG. 31, a rotor 42 having a plurality of blades 46 may be providedas described above. Alternately, a shrouded rotor (e.g., rotor 100 withshroud 102 as shown in FIGS. 10 and 11) may be utilized, as describedabove. Gas 52 at supersonic velocity (Mach>1) is provided to a pluralityof aerodynamic ducts 300. A converging portion 314 is provided using adownstream sidewall 302, which reflects oblique shocks S₁, S₂, etc.generated via leading edge 316. In such embodiment, the radially inwardbounding wall 318 of an aerodynamic duct 300 may be smoothly rounded inconformance with an underlying base, such as a cylinder, or conic shapeas shown in FIGS. 17, 18, and 19, or other smoothly curved shape. In anembodiment, bypass gas outlets 320 may be provided for removal of bypassgas 322 during starting, for example to outboard or inboard sub-chambers(not shown) as otherwise described elsewhere herein. Upon establishmentof normal shock S_(N) at a desired design location for selectedoperational conditions, removal of bypass gas 322 may be terminated. Asgas slows in the subsonic portion (Mach<1) of the aerodynamic ducts 300₁, 300 ₂, 300 ₃, etc., kinetic energy of the gas 52 is converted intogas pressure.

In FIG. 32, yet another embodiment is depicted. Here, a rotor 42 havinga plurality of blades 46 may be provided as described above.Alternately, a shrouded rotor (e.g., rotor 100 with shroud 102 asillustrated in FIGS. 10 and 11) may be utilized. Gas flow 52 atsupersonic velocity (Mach>1) is provided to a plurality of aerodynamicducts 304, here identified in part as individual aerodynamic ducts 304₁, 304 ₂, and 304 ₃. A converging portion 330 is provided using anupstream sidewall 306, which reflects oblique shocks S₁, S₂, etc.generated via leading edge 332. In such embodiment, the radially inwardbounding wall 334 of an aerodynamic duct 304 may be smoothly rounded inconformance with an underlying base, such as a cylinder, or conicalshape as shown in FIGS. 17, 18, and 19, or other smoothly curved shape.In such an embodiment, bypass gas outlets 320 may also be provided forremoval of gas 322 during starting, for example to outboard chambers orinboard sub-chambers (not shown) as otherwise described elsewhereherein. Upon establishment of normal shock S_(N) at a desired designlocation for selected operational conditions, removal of bypass gas 322may be terminated. As gas slows in the subsonic portion (Mach<1) of theaerodynamic ducts 304 ₁, 304 ₂, 304 ₃, etc., kinetic energy of the gasis converted into gas pressure.

In FIG. 33, yet another embodiment using lateral compression, ratherthan radial compression, is depicted. Here, a rotor 42 having aplurality of blades 46 may be provided as described above. Alternately,a shrouded rotor (e.g., rotor 100 with shroud 102 for blades 103 asillustrated in FIGS. 10 and 11) may be utilized. Gas flow 52 atsupersonic velocity (Mach>1) is provided to a plurality of aerodynamicducts 308, here identified in part as individual aerodynamic ducts 308₁, 308 ₂, and 308 ₃. Compression is accomplished in a converging portion340 utilizing both a downstream sidewall 310 and an upstream sidewall312. An upstream leading edge 342 is provided to intercept gas flow 52entering the aerodynamic ducts 308 of a stationary diffuser. A set ofoblique shock waves S₁, S₂, S₃, etc. and a normal shock wave S_(N) aregenerated, and exit gas 344 is provided at subsonic (Mach<1) conditions.In such embodiment, the radially inward bounding wall 334 of anaerodynamic duct 308 ₁, 308 ₂, 308 ₃, etc., may be smoothly rounded inconformance with an underlying base, such as a cylinder, or conic shapeas shown in FIGS. 17, 18, and 19 above, or another shape. In suchembodiment, bypass gas outlets 320 may also be provided for removal ofgas 322 during starting, for example to outboard sub-chambers (notshown) such as described with reference to FIG. 28, 29, or 30, orinboard sub-chambers, for example as otherwise described elsewhereherein with reference to FIGS. 4, 5, and 6.

As shown in FIG. 1, aerodynamic ducts 56 in diffuser 54 may beconstructed with leading edges 350. Certain details pertinent to variousembodiments are shown in FIGS. 8, 8A, 8B, 8C, 15, and 16. In FIG. 15, anembodiment is shown for a stationary diffuser 54 having five (5)aerodynamic ducts 56 ₁ through 56 ₅, and wherein each of suchaerodynamic ducts 56 _(1through5) includes a leading edge 350. In FIG.16, an embodiment is shown for a stationary diffuser 54 having seven (7)aerodynamic ducts 56 ₁ through 56 ₇, and wherein each of suchaerodynamic ducts includes a leading edge 350. Generally, the shaper theleading edge 350, the better performance will be provided, that is,losses will be minimized, when operating at supersonic conditions at theinlet, as compared to use of a leading edge that is not as sharp. In anembodiment, a leading edge 350 may be provided having a leading edgeradius R of from about 0.005 inches to about 0.012 inches, as shown inFIG. 8C. The leading edge 350 may be provided using a sharp leading edgewedge angle theta (θ), which may in an embodiment be between about five(5) degrees and about ten (10) degrees, as shown in FIG. 8A. Also, asseen in FIG. 8B, leading edge 350 may be provided sloping rearward, i.e.in a downstream direction at a slope angle mu (μ) as measured betweenthe leading edge 350 and a tangent line 352 with underlying radiallyinward bounding wall 354. Such sloping leading edge 350 may start at alower front end 356 and end at an upper rear end 358. The leading edge350 may be sealed to or affixed to a radially inward bounding wall 354at the lower front end 356, and may be sealed to or affixed to (forexample, using welded assembly) or otherwise sealingly provided (forexample, machined from a common workpiece) with respect to radiallyoutward bounding wall 360 at the upper rear end 358 of leading edge 350.

Rearward (in the downstream, gas flow direction) from leading edge 350,a partition wall 364 may be utilized. In various embodiments, forexample as seen in FIG. 7, a common partition wall 364 may be utilizedbetween adjacent aerodynamic ducts 56 for example, between individuallyidentified aerodynamic ducts 56 ₁, 56 ₂, etc., through duct 56 ₅ asdepicted in FIGS. 7 and 15. As shown in FIG. 7, partition walls 364 areindividually identified as partition walls 364 ₁, 364 ₂, 364 ₃, etc. asappropriate given the number of aerodynamic ducts 56 utilized. In anembodiment, partition walls 364 may be provided with a thickness T ofabout zero point one zero (0.10) inches, or less. In summary, anefficient compressor may be provided when aerodynamic ducts 56 arelocated adjacent one to another. Such design is even more efficient whenadjacent aerodynamic ducts 56 have a common partition wall 364therebetween. In various embodiments, a leading edge 350 may provide anupstream terminus for a partition wall, such as partition wall 364.

In an embodiment, for example as depicted in FIGS. 1 and 8, a diffuser54 design may include aerodynamic ducts 56 which are polygonal in crosssectional shape, and such shape may include a variety of bounding walls,such as a floor, ceiling, and sidewalls. As used herein, the termradially inward bounding wall has been used to describe what might bealso be considered a floor of an aerodynamic duct. As used herein, theterm radially outward bounding wall has been used to describe what mightbe also considered a ceiling of an aerodynamic duct. As earlier noted,in an embodiment, aerodynamic ducts 56 may have a flow centerline C_(LD)as shown in FIG. 1. Then, in such embodiment, orthogonal to thecenterline line C_(LD), the aerodynamic ducts 56 may be provided havinga parallelogram cross-sectional shape, which may be in an embodiment agenerally rectangular cross sectional shape at various points along theaerodynamic duct 56. In an embodiment, the centerline C_(LD) may begenerally helical. The height H of such a cross-section is shown in FIG.8B, seen radially outward from a radially inward bounding wall 354toward a radially outward bounding wall 360, at an entrance location toan aerodynamic duct 56, namely the lower front end 356 of leading edge350. The width W of such a cross-section is depicted in FIG. 8 asbetween (and within) adjacent partition walls 364 ₁ and 364 ₂. In anembodiment, associated with the just noted cross-sectional shape, theaerodynamic ducts 56 may have an average aspect ratio, expressed aswidth W to height H, of about two to one (2:1), or more. In anembodiment, the aerodynamic ducts 56 may have an average aspect ratio,expressed as width W to height H, of about three to one (3:1), or more.In an embodiment, the aerodynamic ducts 56 may have an average aspectratio, expressed as width W to height H, of about four to one (4:1), ormore.

In various embodiments, the number of aerodynamic ducts 56 may beselected as useful given other design constraints. The number ofaerodynamic ducts 56 included may be one or more, say in the range offrom 1 to 11, or more, for example, 3, 5, 7, 9, or 11 aerodynamic ducts56. The number of aerodynamic ducts for a given design may be selectedas part of a design exercise that takes into account various factorsincluding the direction of gas flow leaving the impulse rotor, and thevelocity provided thereby, and the degree of growth of adverse boundarylayers in configurations of various geometry. In an embodiment, thenumber of leading edges 350 for an inlet in a diffuser 54 may be equalto the number of aerodynamic ducts 56 in a diffuser 54, in a manner assuch parts (e.g., aerodynamic duct 56 ₂) are identified in FIG. 8. Inmany embodiments, design optimization may result in a plurality ofaerodynamic ducts, so that velocity of gas leaving an impulse blade ismaximized and boundary layer growth is minimized. In such embodiments,when optimizing a compressor design, an odd number 3, 5, 7, 9, or 11 ofaerodynamic ducts 56 may be provided, and as just mentioned above, thenumber of leading edges 350 such diffusers 54 would be eleven (11), orless. By selection of an odd number of blades 46 in a rotor 42, an evennumber of aerodynamic ducts 56 may be provided, for example, 2, 4, 6, 8,10, or more. In related parameters, in an exemplary stationary diffuser54, the number of leading edges 350 in a diffuser 54 would be about onehalf (½) or less than the number of blades 46 provided in a rotor 42. Inanother embodiment, the number of leading edges 350 in a diffuser 54would be about one quarter (¼) or less than the number of blades in arotor 42. In a yet more efficient design, it is currently anticipatedthat the number of leading edges 350 in a diffuser 54 would be aboutfifteen percent (15%), or less, of the number of blades in a rotor 42.Minimizing the number of leading edges, and related aerodynamic ducts,minimizes drag and efficiency loss compared to various prior artstators, particularly those utilizing stator blades in numbercommensurate with or equivalent to the number of rotor blades provided.

In addition to improvements in the number, size, and shape of leadingedges 350, and related aerodynamic duct 56 components, the provision ofon-board supersonic shock starting capability, for example by use ofbypass gas passageways, such as bypass gas sub-chambers 114, as shown inFIG. 4 (that is, sub-chambers below the radially inward bounding wall 58of the aerodynamic duct 56) or outboard bypass gas chambers 282 ₁, etc.,as seen in FIG. 30, above a radially outward bounding wall 276) orinternal bypass using internal starting bypass gas passageways 130 asdefined by internal walls 131 of an internal gas passageway housing 133as seen in FIG. 13, provides the ability to design for higher pressureratios in a supersonic compressor. As an example, but not as alimitation, the bypass gas passageways 130 seen in FIG. 13 may beoperable during establishment of a supersonic shock during startup, whenthe compressor 40 shown in FIG. 1 (or compressor 230 of FIG. 9) isdesigned for operating at an inlet relative Mach number of about 1.8,for removal of a quantity of from about eleven percent (11%) by mass toabout nineteen percent (19%) by mass of the selected gas captured at theinlet by an aerodynamic duct 56. As a further example, but not as alimitation, the bypass gas passageways 130 may be operable duringestablishment of a supersonic shock during startup, when a compressor isdesigned for operating at an inlet relative Mach number of about 2.8,for removal of a quantity of from about thirty six percent (36%) by massto about sixty one percent (61%) by mass of the inlet gas captured atthe inlet by an aerodynamic duct 56. Those of skill in the art and towhom this specification is directed will undoubtedly be able tocalculate and thus determine suitable bypass gas quantities that may beuseful or required for enabling aerodynamic ducts used in a particularstator, given compressor design parameters, to swallow an incipientsupersonic shock structure and to thus establish a stable supersonicshock structure at a desired location within the aerodynamic duct(s).Thus, the above noted ranges are to provide to the reader anappreciation of the amount of mass flow that may be required toestablish a stable supersonic shock structure, and thus eliminate anun-started condition in the aerodynamic ducts in a stator. Variousaspects of starting requirements are discussed by Lawlor, in U.S. PatentApplication Publication No. US2009/0196731 A1, Published on Aug. 6,2009, entitled “Method and Apparatus for Starting SupersonicCompressors,” which is incorporated herein in its entirety by thisreference. In particular, FIG. 3 of that publication provides a graphicillustration of typical ranges suitable for starting bypass gas removalrequirements, shown as starting bleed fraction (defined by mass ofbypass gas divided by mass of gas captured by the inlet) fraction, foraerodynamic ducts in a supersonic compressor operating at a selectedinlet relative Mach number.

More generally, a compressor as described herein may be designed forproviding gas to aerodynamic ducts, such as aerodynamic duct 56 ₁ shownin FIG. 1, at an inlet relative Mach number in excess of about 1.8.Further, a compressor as described herein may be designed for an inletrelative Mach number to aerodynamic ducts of at least 2. Even further, acompressor as described herein may be designed for an inlet relativeMach number to aerodynamic ducts of at least 2.5. And, operation ofsupersonic compressors described herein is anticipated to be possible atdesigns having an inlet relative Mach number to aerodynamic ducts inexcess of about 2.5. For many applications, a practical design isanticipated to utilize an inlet relative Mach number to aerodynamicducts between about 2 and about 2.5, inclusive of such boundingparameters. Further, for various applications, as an example and not asa limitation, practical designs may be anticipated to utilize an inletrelative Mach number to aerodynamic ducts in the range of between about2.5 and about 2.8. For other applications, even higher inlet Machnumbers may be practical in various designs, as an example, especiallyfor those gases in which the speed of sound is relatively low, such assome of the refrigerant gases. On the other hand, for applicationshandling gases having a very high speed of sound, such as hydrogen,operation at much lower Mach numbers may provide commercially acceptableresults. Consequently, the Mach number achievable for various designsshould not be considered limited by such above noted suggestions, as anevaluation of design Mach numbers for particular applications mayinclude a variety of design considerations.

Compressors as described herein may be provided for operation within adesign operating envelope having a gas compression ratio of at leastthree (3). In other applications, compressors as described herein may beprovided for operation within a design operating envelope having a gascompression ratio in a stage of compression of at least five (5). In yetother applications, compressors as described herein may be provided foroperation within a design operating envelope having a gas compressionratio in a stage of compression of from about three point seven five(3.75) to about twelve (12). In yet other applications, compressors asdescribed herein may be provided for operation within a design operatingenvelope having a gas compression ratio in a stage of compression offrom about six (6) to about twelve point five (12.5). In certainapplications, compressors as described herein may be provided foroperation within a design operating envelope of gas compression ratiosin a stage of compression of from about twelve (12) to about thirty(30).

When high compression ratios are required by design requirements,multistage compression may be employed, as suggested by theconfiguration schematically depicted for a compressor 400 in FIG. 34. Adriver 402 such as electric motor or other mechanical drive may turn,through gearbox 404 where required, and via shaft 406, a firstcompressor rotor as described herein in a first compression stage 408,to compress entering low pressure gas 410 to provide a dischargeintermediate pressure gas 412. A second compression stage 414, having asecond compressor rotor and stator as described herein, compresses theintermediate pressure gas 412 to provide a high pressure outlet gas 416.In this manner, back to back compression stages may be provided in aplurality of stages as may be desired. Thus, high pressure ratios can beachieved by multistage operation. As an example, but without limitation,such configurations may be provided broadly provide overall pressureratios (in the plurality of stages in series configuration) of fromabout fifty to one (50:1) to about two hundred to one (200:1). Or asanother example, two stages of twenty to one (20:1) each will provide anoverall compression ratio of about four hundred to one (400:1). Finally,it should be noted that multiple stages may also be provided in parallelconfiguration where multiple machines may be desired for capacityconfigurations.

In general, improved supersonic gas compressor designs for compressing aselected gas are provided by the teachings herein. In an embodiment, anexemplary compressor 230 as depicted in FIG. 9 may utilize a casing 234having a low pressure gas inlet 236 and a high pressure gas exit 238. Arotor 100 with shrouded blades 103 may be provided for delivery of aselected gas at supersonic conditions to a stationary diffuser 54 orstator having a plurality of aerodynamic ducts 56, as seen in FIG. 1. Inan embodiment, the aerodynamic ducts 56 may be wrapped helically in adiffuser 54. In an embodiment, adjacent aerodynamic ducts may havecommon partition walls therebetween. The aerodynamic ducts 56 have aconverging portion and a diverging portion that with input of asupersonic gas flow generate a plurality of oblique shock waves (S₁ toS_(x), as seen, for example, in FIGS. 28 through 30) and a normal shockwave (S_(N)) as the selected gas passes through the aerodynamic duct 56.In various designs, aerodynamic ducts 56 may have an inlet relative Machnumber for operation associated with a design operating point selectedwithin a design operating envelope for a selected gas composition, gasquantity, and gas compression ratio. Further, such a compressor mayinclude means for adjusting the effective contraction ratio of some orall of the plurality of aerodynamic ducts, or of each of the aerodynamicducts. The means for adjusting the effective contraction ratio mayinclude bypass gas passageways for discharge of gas 113 from aerodynamicducts to external discharge 118 or recycle 118′ lines as seen in FIG. 4above. The means for adjusting the effective contraction ratio mayinclude internal bypass gas passageways 130, such as using internal gaspassageway housing 133 devices 133 with inlet doors 140 and outlet doors150 as conceptually depicted in FIGS. 13 and 14. The means for adjustingeffective contraction ratio may include geometrically adjustableportions 160 as seen in FIGS. 11 and 12. Even further, as appropriatefor a particular design configuration, means for controlling a boundarylayer of gas flowing through each of the plurality of aerodynamic ductsmay be provided. The means for controlling boundary layers may includeboundary layer outlet bleed ports. The means for controlling boundarylayers may include the use of inlet jets for injection gas into aboundary layer, to energize the same and increase the velocity of theboundary layer to a velocity more closely matching that of bulk fluidflow at a particular location in an aerodynamic duct. The means forcontrolling boundary layers may include the use of one or more vortexgenerators in an aerodynamic duct, to energize a boundary layer bymoving gas via a vortex from a higher velocity bulk flow portion into aslower boundary layer flow, to thereby energize the boundary layer flow.

Various gases or gas mixtures may be selected for compression usingdesigns taught herein. The compression of various hydrocarbon gases,such as ethane, propane, butane, pentane, and hexane, may benefit usingcompressors as taught herein. Further, gases or gas mixtures having amolecular weight of at least that of gaseous nitrogen (MW=28.02) willespecially benefit using the designs taught herein. And, the efficiencyof compression of heavier gases such as carbon dioxide (MW=44.01) may beespecially improved by utilization of the compressor designs as taughtherein. More generally, compression of those gases wherein Mach 1 occursat relatively low velocity, such as that of methane (1440 feet/sec), andlower (such as ammonia, water vapor, air, carbon dioxide, propane,R410a, R22, R134a, R12, R245fa, and R123), may benefit from efficientsupersonic compression.

The compression of various low molecular weight gases, and even thosehaving high sonic velocity, may be efficiently achieved using thedesigns disclosed herein. In some applications, where higher compressionratios are desired, for example but not as a limitation, applicationsinvolving compression ratios in excess of about six (6) or so aresought, useful designs may be provided using the techniques taughtherein. In one example, the compression of hydrogen (MW=2.0158) whichhas a speed of sound of about 4167 feet per second at about 77° F. (1270meters per second at about 25° C.) may be effectively accomplished usingthe compressor configuration(s) taught herein, when constructedutilizing rotors with high strength and using a shrouded bladeconfiguration. Such a design must be able to operate at high rotationalrates to provide sufficient peripheral speed in order achieve a suitablesupersonic design velocity at time of entry of gas to the aerodynamicducts of a stator. As an example, with rotor tip speeds in the range ofabout 2,500 feet per second, using advanced graphite compositeconstruction with shrouded rotor blades, compression ratios of up toabout 5:1 may be achievable using the designs taught herein. Furtheradvances in materials and manufacturing techniques may enable designs ateven higher speeds and pressure ratios, or may provide reduced risk ofmechanical failures when operating at or near the just noted designparameters.

It should be recognized that the stator design taught herein, namelyusing a plurality of aerodynamic ducts, especially when used in ahelical, spiral, helicoidal, or similar curving structure wrapped abouta longitudinal axis, may be especially useful for various statorapplications in supersonic compression, further including, for example,as described in an improved gas turbine apparatus. Thus, the statordesign itself is believed to be a significant improvement in the designof supersonic stators for diffusion of supersonic gas to produce highpressure gas, regardless of application for such gas compression.

Further to the details noted above, it must be reiterated that theaerodynamic ducts described herein may be utilized in configurationsbuilt on various substrate structural designs, and achieve the benefitof high compression ratio operation, while providing necessary featuresfor starting of supersonic operation. In various embodiments, aplurality of aerodynamic ducts may be configured as if wrapped about asurface of revolution, as provided by such static structure. In anembodiment, a suitable static structure may be substantiallycylindrical, and thus, in an embodiment, the ducts may be configuredwrapped around the cylindrical structure. In an embodiment, theaerodynamic ducts of a stationary diffuser may be provided in a spiralconfiguration. In an embodiment the aerodynamic ducts of a stationarydiffuser may be provided in helicoidal configuration, such as may begenerated along a centerline by rotating an entrance plane shape about alongitudinal axis at a fixed rate and simultaneously translating it inthe downstream direction of the longitudinal axis, also at a fixed rate.Thus, the term wrapped around a longitudinal axis shall be considered toinclude wrapping around such various shapes, as applicable.

In summary, the various embodiments using aerodynamic ducts withinternal compression ramps configured as taught herein providesignificantly improved performance over prior art bladed stator designsoperating at supersonic inlet conditions, particularly in their abilityto provide high total and static pressure ratios. In one aspect, this isbecause utilizing a minimum number of aerodynamic ducts, and associatedleading edge structures, reduces loss associated with entry of highvelocity gas into a diffuser. Moreover, the reduced static structurecorrespondingly reduces compressor weight and cost, especially comparedto prior art designs utilizing large numbers of conventional airfoilshaped stator blades.

In the foregoing description, for purposes of explanation, numerousdetails have been set forth in order to provide a thorough understandingof the disclosed exemplary embodiments for the design of a novelsupersonic compressor system for the efficient compression of gases.However, certain of the described details may not be required in orderto provide useful embodiments, or to practice a selected or otherdisclosed embodiments. Further, for descriptive purposes, variousrelative terms may be used. Terms that are relative only to a point ofreference are not meant to be interpreted as absolute limitations, butare instead included in the foregoing description to facilitateunderstanding of the various aspects of the disclosed embodiments. And,various actions or activities in a method described herein may have beendescribed as multiple discrete activities, in turn, in a manner that ismost helpful in understanding the present invention. However, the orderof description should not be construed as to imply that such activitiesare necessarily order dependent. In particular, certain operations maynot necessarily need to be performed precisely in the order ofpresentation. And, in different embodiments of the invention, one ormore activities may be performed simultaneously, or eliminated in partor in whole while other activities may be added. Also, the reader willnote that the phrase “in an embodiment” or “in one embodiment” has beenused repeatedly. This phrase generally does not refer to the sameembodiment; however, it may. Finally, the terms “comprising”, “having”and “including” should be considered synonymous, unless the contextdictates otherwise.

From the foregoing, it can be understood by persons skilled in the artthat a supersonic compressor system has been provided for the efficientcompression of various gases. Although only certain specific embodimentsof the present invention have been shown and described, there is nointent to limit this invention by these embodiments. Rather, theinvention is to be defined by the appended claims and their equivalentswhen taken in combination with the description.

Importantly, the aspects and embodiments described and claimed hereinmay be modified from those shown without materially departing from thenovel teachings and advantages provided, and may be embodied in otherspecific forms without departing from the spirit or essentialcharacteristics thereof. Therefore, the embodiments presented herein areto be considered in all respects as illustrative and not restrictive orlimiting. As such, this disclosure is intended to cover the structuresdescribed herein and not only structural equivalents thereof, but alsoequivalent structures. Numerous modifications and variations arepossible in light of the above teachings. Therefore, the protectionafforded to this invention should be limited only by the claims setforth herein, and the legal equivalents thereof.

1. A compressor, comprising: a rotor having an axis of rotation and aplurality of impulse blades extending into a gas flow passage, saidplurality of impulse blades being sized and shaped to act on a selectedgas to provide a supersonic gas flow; and a diffuser comprising one ormore aerodynamic duct(s) helically arranged about a longitudinal axisand positioned to receive said supersonic gas flow, said one or moreaerodynamic duct(s) comprising (a) a converging portion and a divergingportion, and (b) bypass gas passageways operable to remove at least someof said supersonic gas flow from said converging portion, to therebyadjust an effective contraction ratio in one or more of the said one ormore aerodynamic duct(s); said one or more aerodynamic duct(s) sized andshaped to decelerate said supersonic gas flow to subsonic conditions. 2.A compressor, comprising: a rotor having an axis of rotation and aplurality of impulse blades extending into a gas flow passage, saidplurality of impulse blades sized and shaped to act on a selected gas toprovide a supersonic gas flow; and a diffuser having a longitudinalaxis, comprising one or more aerodynamic duct(s) helically arrangedabout said longitudinal axis and positioned to receive said supersonicgas flow, said one or more aerodynamic duct(s) comprising (a) aconverging portion and a diverging portion, and (b) a geometricallyadjustable portion operable to change the shape and/or location of saidconverging portion, to thereby adjust an effective contraction ratio inone or more of said one or more aerodynamic duct(s); said one or moreaerodynamic duct(s) sized and shaped to decelerate said supersonic gasflow to subsonic conditions.
 3. (canceled)
 4. (canceled)
 5. (canceled)6. A compressor, comprising: a rotor having an axis of rotation and aplurality of impulse blades extending into a gas flow passage, saidplurality of impulse blades sized and shaped to act on a selected gas toprovide a supersonic gas flow; and a diffuser disposed around alongitudinal axis and comprising one or more aerodynamic ducts, said oneor more aerodynamic ducts comprising converging and diverging portions,and having an effective contraction ratio, said one or more aerodynamicducts sized and shaped to decelerate said supersonic gas flow tosubsonic conditions from a selected inlet Mach number, and (a) at leastone of (i) bypass gas passageways or (ii) geometrically adjustableportions operable to adjust said effective contraction ratio, or (iii)both, and (b) boundary layer control structures comprising one or moreof (1) outlet bleed ports for boundary layer removal, (2) inlet jets forenergizing a boundary layer by gas injection, and (3) one or more vortexgenerators.
 7. (canceled)
 8. (canceled)
 9. (canceled)
 10. (canceled) 11.(canceled)
 12. (canceled)
 13. (canceled)
 14. (canceled)
 15. (canceled)16. (canceled)
 17. (canceled)
 18. (canceled)
 19. (canceled) 20.(canceled)
 21. The compressor as set forth in claim 2, or in claim 6,wherein said geometrically adjustable portions are positionable betweenan open, startup condition wherein said converging portion allowssufficient flow of said selected gas through said one or moreaerodynamic ducts to establish and position a normal shock within saidone or more aerodynamic ducts, and a closed, operating condition inwhich said converging portion is set to a selected operating position.22. (canceled)
 23. (canceled)
 24. (canceled)
 25. The compressor as setforth in claim 6, wherein said inlet jets are oriented to inject gasinto a boundary layer in a flow of said selected gas in said one or moreaerodynamic ducts.
 26. The compressor as set forth in claim 6, furthercomprising inlet ports and injection gas chambers, said injection gaschambers adjacent said one or more aerodynamic ducts, said injection gaschambers in fluid communication with said inlet ports, said injectiongas chambers configured for passage therethrough of said selected gasfor injection via said inlet ports.
 27. The compressor as set forth inclaim 25, wherein said inlet jets are sized and shaped to provide a gasjet that increases momentum of said flow of said selected gas.
 28. Thecompressor as set forth in claim 6, wherein said boundary layer controlstructures are configured as said one or more vortex generators.
 29. Thecompressor as set forth in claim 28, wherein said one or more vortexgenerators are located in said converging portion.
 30. The compressor asset forth in claim 28, wherein said one or more vortex generators arelocated in said diverging portion.
 31. The compressor as set forth inclaim 28, wherein each of said one or more vortex generators comprise abase with a forward end and a leading edge extending outward to anoutward end.
 32. The compressor as set forth in claim 31, wherein saidleading edge comprises a discontinuity along said leading edge, forgenerating a single vortex.
 33. The compressor as set forth in claim 31,wherein said one or more vortex generators comprises at least two vortexgenerators, and wherein each of said at least two vortex generatorsgenerate a single vortex configured to control boundary layer in theaerodynamic duct.
 34. The compressor as set forth in claim 31, whereinsaid outward end is located at a height H_(Z) above said base.
 35. Thecompressor as set forth in claim 34, wherein said aerodynamic duct has aheight H_(D), and wherein said height H_(Z) above said base is aboutfifty percent (50%) or less, of said height H_(D).
 36. The compressor asset forth in claim 34, wherein said aerodynamic duct has a height H_(D),and wherein said height H_(Z) above said base is about twenty fivepercent (25%), or less, of said height H_(D).
 37. The compressor as setforth in claim 6, wherein a plurality of vortex generators are providedin each of said aerodynamic ducts.
 38. The compressor as set forth inclaim 6, wherein one or more of said one or more aerodynamic ducts arehelically arranged about said longitudinal axis.
 39. The compressor asset forth in claim 38, wherein said one or more of said aerodynamicducts are helically arranged at a substantially constant helical angleabout said longitudinal axis.
 40. (canceled)
 41. A supersonic gascompressor for compressing a selected gas, comprising: a casingcomprising a low pressure gas inlet and a high pressure gas exit; arotor comprising a plurality of blades and configured to act on aselected gas to impart axial and tangential velocity thereto to providea supersonic gas flow; a stator comprising a diffuser including one ormore aerodynamic ducts configured for diffusing a gas received therein,said one or more aerodynamic ducts each having a converging portion, adiverging portion, and an effective contraction ratio, such that, withinput of a supersonic gas flow, each aerodynamic duct generates aplurality of oblique shock waves (S₁ to S_(x)) and a normal shock wave(S_(N)) in said selected gas as said selected gas passes therethrough,said one or more aerodynamic ducts having an inlet relative Mach numberfor operation associated with a design operating point selected within adesign operating envelope for a selected gas composition, gas quantity,and gas compression ratio, said one or more aerodynamic ducts comprising(a) bypass gas passageways or a geometrically adjustable portion, orboth, operable to adjust said effective contraction ratio, and (b)boundary layer control structures, said boundary layer controlstructures comprising one or more of (1) outlet bleed ports for boundarylayer removal, (2) inlet jets for energizing a boundary layer by gasinjection, and (3) one or more vortex generators.
 42. The compressor asset forth in claim 41, wherein said one or more aerodynamic ducts arehelically arranged around a longitudinal axis.
 43. (canceled) 44.(canceled)
 45. (canceled)
 46. (canceled)
 47. (canceled)
 48. (canceled)49. (canceled)
 50. (canceled)
 51. (canceled)
 52. (canceled) 53.(canceled)
 54. (canceled)
 55. (canceled)
 56. (canceled)
 57. (canceled)58. (canceled)
 59. (canceled)
 60. (canceled)
 61. (canceled) 62.(canceled)
 63. (canceled)
 64. (canceled)
 65. (canceled)
 66. (canceled)67. (canceled)
 68. (canceled)
 69. A supersonic gas compressor forcompressing a selected gas, comprising: a casing comprising a lowpressure gas inlet and a high pressure gas exit; an impulse rotorconfigured to act on a selected gas to impart axial and tangentialvelocity thereto to provide a supersonic gas flow; a stator comprising:a diffuser including a plurality of aerodynamic ducts configured fordiffusing a gas received therein, said plurality of aerodynamic ductshelically arranged in adjacent position, and having a converging portionand a diverging portion that, with input of said supersonic gas flow,generate a plurality of oblique shock waves (S₁ to S_(x)) and a normalshock wave (S_(N)) in said selected gas as said selected gas passesthrough said aerodynamic ducts, said plurality of aerodynamic ductshaving an inlet relative Mach number for operation associated with adesign operating point selected within a design operating envelope for aselected gas composition, gas quantity, effective contraction ratio, andgas compression ratio, and said plurality of aerodynamic ducts furthercomprising means for adjusting said effective contraction ratio of someor all of said plurality of aerodynamic ducts, and means for controllinga boundary layer of gas flowing through said plurality of aerodynamicducts.
 70. (canceled)
 71. The compressor as set forth in claim 69,wherein said means for adjusting the effective contraction ratiocomprises geometrically adjustable portions in said plurality ofaerodynamic ducts, said geometrically adjustable portions positionablebetween an open, startup condition wherein said converging portionallows increased flow of said selected gas through said plurality ofaerodynamic ducts, and a closed, operating condition in which saidconverging portion is set to a selected operating position.
 72. Thecompressor as set forth in claim 69, wherein means for controlling aboundary layer of gas flowing through said plurality of aerodynamicducts comprises inlet jets.
 73. The compressor as set forth in claim 69,wherein means for controlling a boundary layer of gas flowing throughsaid plurality of aerodynamic ducts comprises boundary layer outletbleed ports.
 74. The compressor as set forth in claim 69, wherein meansfor controlling a boundary layer of gas flowing through said pluralityof aerodynamic ducts comprises one or more vortex generators.
 75. Thecompressor as set forth in claim 41, or in claim 69, wherein said designoperating envelope comprises at least one stage having a gas compressionratio of at least
 3. 76. The compressor as set forth in claim 41, or inclaim 69, wherein said design operating envelope comprises at least onestage having a gas compression ratio of at least
 5. 77. The compressoras set forth in claim 41, or in claim 69, wherein said design operatingenvelope comprises at least one stage having a gas compression ratio offrom about 6 to about 12.5.
 78. The compressor as set forth in claim 41,or in claim 69, wherein said design operating envelope comprises atleast one stage having a gas compression ratio of from about 12 to about30.
 79. (canceled)
 80. (canceled)
 81. (canceled)
 82. (canceled) 83.(canceled)
 84. (canceled)
 85. (canceled)
 86. The compressor as set forthin claim 1, wherein said one or more aerodynamic ducts comprise boundingwalls, and further comprising outlet bleed ports in one or more of saidbounding walls.
 87. The compressor as set forth in claim 1, furthercomprising inlet jets configured to energize a boundary layer by gasinjection in said one or more aerodynamic ducts.
 88. The compressor asset forth in claim 1, further comprising one or more vortex generatorsin said one or more aerodynamic ducts configured to energize a boundarylayer.
 89. The compressor as set forth in claim 2, or in claim 6,wherein said rotor further comprises a shroud for said plurality ofimpulse blades.
 90. The compressor as set forth in claim 2, or in claim6, wherein said rotor is effectively sealed with said diffuser, so as tominimize gas leakage during flow therebetween.
 91. The compressor as setforth in claim 6, wherein said selected gas passing through said rotoris turned by an angle alpha (α) of at least ninety (90) degrees.
 92. Thecompressor as set forth in claim 6, wherein said selected gas passingthrough said rotor is turned by an angle alpha (α) of at least onehundred (100) degrees.
 93. The compressor as set forth in claim 6,wherein said selected gas passing through said rotor is turned by anangle alpha (α) of at least one hundred ten (110) degrees.
 94. Thecompressor as set forth in claim 6, wherein said selected gas passingthrough said rotor is turned by an angle alpha (α) of between aboutninety (90) degrees and about one hundred sixty (160) degrees.
 95. Thecompressor as set forth in claim 6, wherein said selected gas passingthrough said rotor is turned by an angle alpha (α) of between about onehundred (112) degrees and about one hundred fourteen (114) degrees. 96.The compressor as set forth in claim 2, or in claim 6, wherein each ofsaid plurality of blades has a hub end, a tip end, and a trailing edge,and said supersonic gas flow is provided at said trailing edge of eachof said plurality of blades from said hub end to said tip end.
 97. Thecompressor as set forth in claim 6, wherein said geometricallyadjustable portions, by change in position, change the contraction ratioof one or more of said one or more aerodynamic ducts.
 98. The compressoras set forth in claim 97, wherein said geometrically adjustable portionsfurther comprise pivotable members and actuators, said pivotable membersdriven by said actuators, and wherein said geometrically adjustableportions are sized and shaped to change the shape of said convergingportion of said one or more of said one or more aerodynamic ducts whensaid geometrically adjustable portions are moved with said actuators.99. The compressor as set forth in claim 6, wherein said diffusercomprises a stationary diffuser.
 100. The compressor as set forth inclaim 6, or in claim 41, wherein each aerodynamic duct of said one ormore aerodynamic ducts comprises a leading edge associated therewith.101. The compressor as set forth in claim 100, wherein said leading edgecomprises a leading edge radius of from about 0.005 inches to about0.012 inches.
 102. The compressor as set forth in claim 100, whereinsaid leading edge defines a leading edge wedge angle of between aboutfive (5) degrees and about ten (10) degrees.
 103. The compressor as setforth in claim 100, further comprising a partition wall downstream fromsaid leading edge.
 104. The compressor as set forth in claim 103,wherein said partition wall divides adjacent aerodynamic ducts, andwherein said leading edge comprises an upstream terminus of saidpartition wall.
 105. The compressor as set forth in claim 103, whereinsaid partition wall has a thickness T of about 0.100 inches, or less.106. The compressor as set forth in claim 100, wherein said plurality ofblades comprises a number B of blades, and wherein a number N of saidone or more aerodynamic ducts are provided, and wherein B and N areselected to avoid harmonic interference between said plurality of bladesand said one or more aerodynamic ducts.
 107. The compressor as set forthin claim 100, wherein each of said one or more aerodynamic ducts has acenterline, and wherein orthogonal to said centerline, one or more ofsaid one or more aerodynamic ducts have a generally parallelogramcross-sectional shape.
 108. The compressor as set forth in claim 107,wherein associated with said cross-sectional shape, said one or moreaerodynamic ducts have an average aspect ratio, expressed as width toheight, of about two to one (2:1), or more.
 109. The compressor as setforth in claim 107, wherein associated with said cross-sectional shape,said one or more of aerodynamic ducts have an average aspect ratio,expressed as width to height, of about three to one (3:1), or more. 110.The compressor as set forth in claim 107, wherein associated with saidcross-sectional shape, said one or more aerodynamic ducts have anaverage aspect ratio, expressed as width to height, of about four to one(4:1), or more.
 111. The compressor as set forth in claim 100, whereinthe number of leading edges in said diffuser is eleven (11), or less.112. The compressor as set forth in claim 100, wherein the number ofleading edges in said diffuser is about one half (½) or less than thenumber of blades B in said rotor.
 113. The compressor as set forth inclaim 100, wherein the number of leading edges in said diffuser is aboutone quarter (¼) or less than the number of blades B in said rotor. 114.The compressor as set forth in claim 100, wherein the number of leadingedges in said diffuser is about fifteen percent (15%), or less, of thenumber of blades B in said rotor.
 115. The compressor as set forth inclaim 6, or in claim 41, wherein said inlet relative Mach number of saidone or more aerodynamic ducts is in excess of 1.5.
 116. The compressoras set forth in claim 6, or in claim 41, wherein said inlet relativeMach number of said one or more aerodynamic ducts is in excess of 1.8.117. The compressor as set forth in claim 6, or in claim 41, whereinsaid inlet relative Mach number of said one or more aerodynamic ducts isat least
 2. 118. The compressor as set forth in claim 6, or in claim 41,wherein said inlet relative Mach number of said one or more aerodynamicducts is at least 2.5.
 119. The compressor as set forth in claim 6, orin claim 41, wherein said inlet relative Mach number of said one or moreaerodynamic ducts is in excess of about 2.5.
 120. The compressor as setforth in claim 6, or in claim 41, wherein said inlet relative Machnumber of said one or more aerodynamic duct is between about 2 and about2.5.
 121. The compressor as set forth in claim 6, or in claim 41,wherein said inlet relative Mach number of said one or more aerodynamicduct is between about 2.5 and about 2.8.
 122. The compressor as setforth in claim 6, or in claim 41, wherein said one or more aerodynamicducts are located adjacent one to another.
 123. The compressor as setforth in claim 122, wherein said adjacent aerodynamic ducts have acommon partition wall therebetween.
 124. The compressor as set forth inclaim 6, or in claim 41, or in claim 69, wherein said selected gascomprises one or more hydrocarbon gases.
 125. The compressor as setforth in claim 6, or in claim 41, or in claim 69, wherein said selectedgas comprises a gas having a molecular weight of at least that ofnitrogen.
 126. The compressor as set forth in claim 125, wherein saidselected gas comprises carbon dioxide.
 127. The compressor as set forthin claim 41, or in claim 69, wherein said selected gas compriseshydrogen.
 128. The compressor as set forth in claim 41, or in claim 69,wherein compression in said plurality of aerodynamic ducts isaccomplished in a channel between spaced apart sidewalls.
 129. Thecompressor as set forth in claim 6, or in claim 41, or in claim 69,wherein compression in said plurality of aerodynamic ducts isaccomplished between radially spaced apart bounding surfaces.
 130. Thecompressor as set forth in claim 6, or in claim 42, wherein said one ormore aerodynamic ducts are helically arranged at a helical angle psi (ψ)in a range of from about forty-five degrees (45°) to about eightydegrees (80°).